Sample records for solid propellant motors

A case bonded end burning solidpropellant rocket motor is described. A propellant with sufficiently low modulus to avoid chamber buckling on cooling from cure and sufficiently high elongation to sustain the stresses induced without cracking is used. The propellant is zone cured within the motor case at high pressures equal to or approaching the pressure at which the motor will operate during combustion. A solidpropellantmotor with a burning time long enough that its spacecraft would be limited to a maximum acceleration of less than 1 g is provided by one version of the case bonded end burning solidpropellantmotor of the invention.

The characteristics of a solidpropellant rocket engine with a controlled rate of thrust buildup to a desired thrust level are discussed. The engine uses a regressive burning controlled flow solidpropellant igniter and a progressive burning main solidpropellant charge. The igniter is capable of operating in a vacuum and sustains the burning of the propellant below its normal combustion limit until the burning propellant surface and combustion chamber pressure have increased sufficiently to provide a stable chamber pressure.

Typical SolidPropellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

Typical SolidPropellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

The basic principles underlying propulsion by rocket motor are examined together with the configuration of a solidpropellantmotor. Solidpropellants and their preparation are discussed, giving attention to homogeneous propellants, composite propellants, energetic considerations in choosing a solidpropellant, the processing of composite propellants, and some examples of new developments. The performance of solidpropellants is investigated, taking into account characteristics velocity, the specific impulse, and performance calculations. Aspects of propellant development considered include nonperformance requirements for solidpropellants, the approach to development, propellant mechanical properties, and future trends.

In a study conducted to evaluate flow parameters of uncured solid composite propellants during motor casting, two motors (1.8M-lb grain wt) were cast with a PBAN propellant exhibiting good flow characteristics in a 260-in. dia solid rocket motor. Attention is given to the effects of propellant compositional and processing variables on apparent viscosity as they pertain to rheological behavior and grain defect formation during casting. It is noted that optimized flow behavior is impaired with solidpropellant loading. Non-Newtonian pseudoplastic flow is observed, which is dependent upon applied shear stress and the age of the uncured propellant.

The properties of a propellant developed for the NASA Advanced Solid Rocket Motor (ASRM) are described in terms of its composition, performance, and compliance to NASA specifications. The class 1.3 HTPB/AP/A1 propellant employs an ester plasticizer and the content of ballistic solids is set at 88 percent. Ammonia evolution is prevented by the utilization of a neutral bonding agent which allows continuous mixing. The propellant also comprises a bimodal AP blend with one ground fraction, ground AP of at least 20 microns, and ferric oxide to control the burning rate. The propellant's characteristics are discussed in terms of tradeoffs in AP particle size and the types of Al powder, bonding agent, and HTPB polymer. The size and shape of the ballistic solids affect the processability, ballistic properties, and structural properties of the propellant. The revised baseline composition is based on maximizing the robustness of in-process viscosity, structural integrity, and burning-rate tailoring range.

The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

Full Text Available The special conditions of employment of commercial rockets in the sea environment has opened up new possibilities of improving motor performance. The interesting method suggests supplying water into the running motor. This paper reports the calculations and experiments carried out with solidpropellant model setups. The results prove the validity of the proposed method and allow the refinement of calculation techniques for the prediction of solid rocket motor performance characteristics. The serviceability of the solidpropellant charges working in combination with water is demonstrated. A mathematical model is proposed for the operation of a hydrocombined propellantmotor with water and powdered additives applied to the combustion chamber."

A miniature solidpropellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solidpropellantmotors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.

As part of the Shuttle Exhaust Effects Panel (SEEP) program for fiscal year 1973, a limited study was performed to determine the feasibility of minimizing the environmental impact associated with the operation of the solid rocket booster motors (SRBMs) in projected space shuttle launches. Eleven hypothetical and two existing limited-experience propellants were evaluated as possible alternates to a well-proven state-of-the-art reference propellant with respect to reducing emissions of primary concern: namely, hydrogen chloride (HCl) and aluminum oxide (Al2O3). The study showed that it would be possible to develop a new propellant to effect a considerable reduction of HCl or Al2O3 emissions. At the one extreme, a 23% reduction of HCl is possible along with a ll% reduction in Al2O3, whereas, at the other extreme, a 75% reduction of Al2O3 is possible, but with a resultant 5% increase in HCl.

Results of research aimed at improving the predictability of off nominal internal ballistics performance of solidpropellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.

A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solidpropellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solidpropellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

A comprehensive numerical analysis has been carried out to study the detailed physical and chemical processes involved in the combustion of homogeneous propellant in a rocket motor. The formulation is based on the time-dependent full Navier-Stokes equations, with special attention devoted to the chemical reactions in both gas and condensed phases. The turbulence closure is achieved using both the Baldwin-Lomax algebraic model and a modified k-epsilon two-equation scheme with a low Reynolds number and near-wall treatment. The effects of variable thermodynamic and transport properties are also included. The system of governing equations are solved using a multi-stage Runge-Kutta shceme with the source terms treated implicitly. Preliminary results clearly demonstrate the presence of various combustion regimes in the vicinity of propellant surface. The effects of propellant combustion on the motor internal flowfields are investigated in detail.

Set of gauges on solid-propellant rocket motor with electrically insulating case measures advance of combustion front and local erosion rates of propellant and insulation. Data furnished by gauges aid in motor design, failure analysis, and performance prediction. Technique useful in determining propellant uniformity and electrical properties of exhaust plum. Gauges used both in flight and on ground. Foilgauge technique also useful in basic research on pulsed plasmas or combustion of solids.

Reports are presented at the meeting at the University of Arizona on the study of predictable and reliable solid rocket motors. The following subject areas were covered: present state and trends in the research of solidpropellants; the University of Arizona program in solidpropellants, particularly in mixing (experimental and analytical results are presented).

Presented here is a review of the experimental and modeling work concerning erosive burning of solidpropellants (augmentation of burning rate by flow of product gases across a burning surface). A brief introduction describes the motor design problems caused by this phenomenon, particularly for low port/throat area ratio motors and nozzleless motors. Various experimental techniques for measuring crossflow sensitivity of solidpropellant burning rates are described, with the conclusion that accurate simulation of the flow, including upstream flow development, in actual motors is important since the degree of erosive burning depends not only on local mean crossflow velocity and propellant nature, but also upon this upstream development. In the modeling area, a brief review of simplified models and correlating equations is presented, followed by a description of more complex numerical analysis models. Both composite and double-base propellant models are reviewed. A second generation composite model is shown to give good agreement with data obtained in a series of tests in which composite propellant composition and heterogeneity (particle size distribution) were systematically varied. Finally, the use of numerical models for the development of erosive burning correlations is described, and a brief discussion of scaling is presented.

A comprehensive analytical model which considers time and space development of the flow field in solidpropellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

Hercules' clean propellant development research is exploring three major types of clean propellant: (1) chloride-free formulations (no chlorine containing ingredients), being developed on the Clean Propellant Development and Demonstration (CPDD) contract sponsored by Phillips Laboratory, Edwards Air Force Base, CA; (2) low HCl scavenged formulations (HCl-scavenger added to propellant oxidized with ammonium perchlorate (AP)); and (3) low HCl formulations oxidized with a combination of AN and AP (with or without an HCl scavenger) to provide a significant reduction (relative to current solid rocket boosters) in exhaust HCl. These propellants provide performance approaching that of current systems, with less than 2 percent HCl in the exhaust, a significant reduction (greater than or equal to 70 percent) in exhaust HCl levels. Excellent processing, safety, and mechanical properties were achieved using only readily available, low cost ingredients. Two formulations, a sodium nitrate (NaNO3) scavenged HTPB and a chloride-free hydroxy terminated polyether (HTPE) propellant, were characterized for ballistic, mechanical, and rheological properties. In addition, the hazards properties were demonstrated to provide two families of class 1.3, 'zero-card' propellants. Further characterization is planned which includes demonstration of ballistic tailorability in subscale (one to 70 pound) motors over the range of burn rates required for retrofit into current Hercules space booster designs (Titan 4 SRMU and Delta 2 GEM).

Multi-phase flow field simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM)is in- vestigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and form liquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flow simulations have been performed on a large scale motor,to observe the effect of the flowfield on the chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an expo- nential relationship exists between the droplet diameter,d and mass fraction of droplets with diameter greater than d have been used for the simulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-1 00μm are used,as being the most common droplets.In this approach the complete range of particle sizes is divided into a set of discrete size ranges,each to be defined by single stream that is part of the group.Roe scheme-flux differencing splitting based on approximate Riemann problem has been used to simulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect, Spalart-Allmaras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRM flow dynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the parti- cles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

A total of 68 quench tests were conducted in a vented bomb assembly (VBA). Designed to simulate full-scale motor operating conditions, this laboratory apparatus uses a 2-inch-diameter, end-burning propellant charge and an insulated disc of consolidated hydrated aluminum sulfate along with the explosive charge necessary to disperse the salt and inject it onto the burning surface. The VBA was constructed to permit variation of motor design parameters of interest; i.e., weight of salt per unit burning surface area, weight of explosive per unit weight of salt, distance from salt surface to burning surface, incidence angle of salt injection, chamber pressure, and burn time. Completely satisfactory salt quenching, without re-ignition, occurred in only two VBA tests. These were accomplished with a quench charge ratio (QCR) of 0.023 lb salt per square inch of burning surface at dispersing charge ratios (DCR) of 13 and 28 lb of salt per lb of explosive. Candidate materials for insulating salt charges from the rocket combustion environment were evaluated in firings of 5-inch-diameter, uncured end-burner motors. A pressed, alumina ceramic fiber material was selected for further evaluation and use in the final demonstration motor.

One United Technology Center FW-4S solid-propellant rocket motor was fired at an average simulated altitude of 103,000 ft while spinning about its axial centerline at 180 rpm. The objectives of the test program were to determine motor altitude ballistic performance including the measurement of the nonaxial thrust vector and to demonstrate structural integrity of the motor case and nozzle. These objectives are presented and discussed.

The Monte Carlo method was used to investigate thrust imbalance and its first time derivative throughtout the burning time of pairs of solid rocket motors firing in parallel. Results obtained compare favorably with Titan 3 C flight performance data. Statistical correlations of the thrust imbalance at various times with corresponding nominal trace slopes suggest several alternative methods of predicting thrust imbalance. The effect of circular-perforated grain deformation on internal ballistics is discussed, and a modified design analysis computer program which permits such an evaluation is presented. Comparisons with SRM firings indicate that grain deformation may account for a portion of the so-called scale factor on burning rate between large motors and strand burners or small ballistic test motors. Thermoelastic effects on burning rate are also investigated. Burning surface temperature is calculated by coupling the solid phase energy equation containing a strain rate term with a model of gas phase combustion zone using the Zeldovich-Novozhilov technique. Comparisons of solutions with and without the strain rate term indicate a small but possibly significant effect of the thermoelastic coupling.

This research aims to reduce the cost of the solid rocket motor production, mainly solidpropellant. The production process of the solid rocket propellant are usually employed the multi-batch mixing. However, this study using a peristaltic pump as a mixer will lead to the continuous process. The pump system can mix the powder materials for propellant and we consider that it will make the slurry of the solidpropellant efficiently by the mechanism of the fluid dynamics in the pump.

The Marshall Space Flight Center (MSFC) engineers test fired a 26-foot long, 100,000-pound-thrust solid rocket motor for 30 seconds at the MSFC east test area, the first test firing of the Modified NASA Motor (M-NASA Motor). The M-NASA Motor was fired in a newly constructed stand. The motor is 48-inches in diameter and was loaded with two propellant cartridges weighing a total of approximately 12,000 pounds. The purpose of the test was to learn more about solid rocket motor insulation and nozzle materials and to provide young engineers additional hands-on expertise in solid rocket motor technology. The test is a part of NASA's Solid Propulsion Integrity Program, that is to provide NASA engineers with the techniques, engineering tools, and computer programs to be able to better design, build, and verify solid rocket motors.

Asbestos has been used for many years as an ablation inhibitor in insulating materials. It has been a constituent of the AS/NBR insulation used to protect the steel case of the RSRM (Reusable Solid Rocket Motor) since its inception. This paper discusses the development of a potential replacement RSRM insulation design, several of the numerous design issues that were worked and processing problems that were resolved. The earlier design demonstration on FSM-5 (Flight Support Motor) of the selected 7% and 11% Kevlar(registered) filled EPDM (KF/EPDM) candidate materials was expanded. Full-scale process simulation articles were built and FSM-8 was manufactured using multiple Asbestos Free (AF) components and materials. Two major problems had to be overcome in developing the AF design. First, bondline corrosion, which occurred in the double-cured region of the aft dome, had to be eliminated. Second, KF/EPDM creates high levels of electrostatic energy (ESE), which does not readily dissipate from the insulation surface. An uncontrolled electrostatic discharge (ESD) of this surface energy during many phases of production could create serious safety hazards. Numerous processing changes were implemented and a conductive paint was developed to prevent exposed external insulation surfaces from generating ESE/ESD. Additionally, special internal instrumentation was incorporated into FSM-8 to record real-time internal motor environment data. These data included inhibitor insulation erosion rates and internal thermal environments. The FSM-8 static test was successfully conducted in February 2000 and much valuable data were obtained to characterize the AF insulation design.

The use of holography, high speed motion pictures, light scattering measurements, and post-fire particle collection/scanning electron microscopic examination to study the combustion of composite solidpropellants is discussed. The relative advantages and disadvantages of the different experimental techniques for obtaining two-phase flow characteristics within the combustion environment of a solidpropellant grain are evaluated. Combustion bomb studies using high speed motion pictures and post-fire residue analysis were completed for six low metal content propellants. Resolution capabilities and the relationships between post-fire residue and motion picture data are determined. Initial testing using a holocamera together with a 2D windowed motor is also described.

Two different independent variable forms, a difference form and a ratio form, were investigated for correlating the normalized magnitude of the measured erosive burning rate augmentation above the threshold in terms of the amount that the driving parameter (mass flux or Reynolds number) exceeds the threshold value for erosive augmentation at the test condition. The latter was calculated from the previously determined threshold correlation. Either variable form provided a correlation for each of the two motor size data bases individually. However, the data showed a motor size effect, supporting the general observation that the magnitude of erosive burning rate augmentation is reduced for larger rocket motors. For both independent variable forms, the required motor size scaling was attained by including the motor port radius raised to a power in the independent parameter. A boundary layer theory analysis confirmed the experimental finding, but showed that the magnitude of the scale effect is itself dependent upon scale, tending to diminish with increasing motor size.

A new nonlinear optimal and explicit guidance law is presented in this paper for launch vehicles propelled by solidmotors. It can ensure very high terminal precision despite not having the exact knowledge of the thrust-time curve apriori. This was motivated from using it for a carrier launch vehicle in a hypersonic mission, which demands an extremely narrow terminal accuracy window for the launch vehicle for successful initiation of operation of the hypersonic vehicle. The proposed explicit guidance scheme, which computes the optimal guidance command online, ensures the required stringent final conditions with high precision at the injection point. A key feature of the proposed guidance law is an innovative extension of the recently developed model predictive static programming guidance with flexible final time. A penalty function approach is also followed to meet the input and output inequality constraints throughout the vehicle trajectory. In this paper, the guidance law has been successfully validated from nonlinear six degree-of-freedom simulation studies by designing an inner-loop autopilot as well, which enhances confidence of its usefulness significantly. In addition to excellent nominal results, the proposed guidance has been found to have good robustness for perturbed cases as well.

This photograph shows the SolidPropellant Test Article (SPTA) test stand with the Modified Nasa Motor (M-NASA) test article at the Marshall Space Flight Center (MSFC). The SPTA test stand, 12-feet wide by 12-feet long by 24-feet high, was built in 1989 to provide comparative performance data on nozzle and case insulation material and to verify thermostructural analysis models. A modified NASA 48-inch solidmotor (M-NASA motor) with a 12-foot blast tube and 10-inch throat makes up the SPTA. The M-NASA motor is being used to evaluate solid rocket motor internal non-asbestos insulation materials, nozzle designs, materials, and new inspection techniques. New internal motor case instrumentation techniques are also being evaluated.

400 jm to reach the maximum flame temperature, a distance that can be reduced by replacing the HTPB binder with a polyester or CMDB binder. The...the dark zone for propellants similar to HIX2 is 2-2.5 mm at 1.8 MPa (18 atm, 265 psia) (Ref. 22,187). In contrast, the dark zone for HMX CMDB ...propellants eliminates the dark zone is not surprising, since TMETN is a nitrate ester as was the double-base matrix of Kubota’s HMX CMDB propellant. A

It is pointed out that in-process tests to verify quality and detect discrepant propellant which could compromise motor performance are essential elements of the solid composite propellant manufacturing process. The successful performance of the 260SL-1 and 260SL-2 motors aptly verified the controls used for manufacturing the propellant. The present investigation is concerned with the selected control parameters, and their relationships to composition and final propellant properties. Control performance is evaluated by comparison with processing data experienced in the manufacture of the propellant for the 260SL-1 motor. It is found that the in-process quality verification controls utilized in the propellant manufacturing process for the 260-in. diameter motor contributed significantly to the confidence of successful and predictable motor performance.

It is pointed out that in-process tests to verify quality and detect discrepant propellant which could compromise motor performance are essential elements of the solid composite propellant manufacturing process. The successful performance of the 260SL-1 and 260SL-2 motors aptly verified the controls used for manufacturing the propellant. The present investigation is concerned with the selected control parameters, and their relationships to composition and final propellant properties. Control performance is evaluated by comparison with processing data experienced in the manufacture of the propellant for the 260SL-1 motor. It is found that the in-process quality verification controls utilized in the propellant manufacturing process for the 260-in. diameter motor contributed significantly to the confidence of successful and predictable motor performance.

The potential environmental effects of the exhaust products of conventional rocket propellants have been assessed by various groups. Areas of concern have included stratospheric ozone, acid rain, toxicity, air quality and global warming. Some of the studies which have been performed on this subject have concluded that while the impacts of rocket use are extremely small, there are propellant development options which have the potential to reduce those impacts even further. This paper discusses the various solidpropellant options which have been proposed as being more environmentally benign than current systems by reducing HCI emissions. These options include acid neutralized, acid scavenged, and nonchlorine propellants. An assessment of the acid reducing potential and the viability of each of these options is made, based on current information. Such an assessment is needed in order to judge whether the potential improvements justify the expenditures of developing the new propellant systems.

Rapid evolution in the structure of military forces worldwide is resulting in the retirement of numerous weapon systems. Many of these systems include rocket motors containing highly energetic propellants based on hazardous nitrocellulose/nitroglycerin (NC/NG) mixtures. Even as the surplus quantities of such material increases, however, current disposal methods -- principally open burning and open detonation (OB/OD) -- are coming under close scrutiny from environmental regulators. Environmentally conscious alternatives to disposal of propellant and explosives are thus receiving renewed interest. Recycle and reuse alternatives to OB/OD appear particularly attractive because some of the energetic materials in the inventories of surplus weapon systems represent potentially valuable resources to the commercial explosives and chemical industries. The ability to reclaim such resources is therefore likely to be a key requirement of any successful technology of the future in rocket motor demilitarization. This document consists of view graphs from the poster session.

Several studies are described of the chemistry of solidpropellant combustion which employed a fast-scanning optical spectrometer. Expanded abstracts are presented for four of the studies which were previously reported. One study of the ignition of composite propellants yielded data which suggested early ammonium perchlorate decomposition and reaction. The results of a study of the spatial distribution of molecular species in flames from uncatalyzed and copper or lead catalyzed double-based propellants support previously published conclusions concerning the site of action of these metal catalysts. A study of the ammonium-perchlorate-polymeric-fuel-binder reaction in thin films, made by use of infrared absorption spectrometry, yielded a characterization of a rapid condensed-phase reaction which is likely important during the ignition transient and the burning process.

The structural properties of solidpropellant rocket grains were studied to determine the propellant resistance to stresses. Grain geometry, thermal properties, mechanical properties, and failure modes are discussed along with design criteria and recommended practices.

Objective To investigate the safety of solid rocket motor(SRM)when it was cook-off. Methods Finite element model of solid SRM was established, and the temperature distribution and the explosion delay time of propellant in fast cook-off mode and slow cook- off mode were computed. Results Propellant reached its critical temperature (352 ℃) after 47 h slow cook-off, while it reached its critical temperature (355 ℃) after 697 s fast cook-off. Conclusion It was proven that the thermal diffusivity in fast cook-off mode was greater than that in slow cook-off mode, while the temperature gradient had an opposite trend. The reaction position of propellant was different in the two different working modes when it reached critical temperature, and the thermal storage capacity of propellant was dependent on its thickness.%目的：研究固体火箭发动机遭受火烤时的安全性。方法建立发动机有限元模型，计算推进剂在慢速烤燃和快速烤燃工况下的温度分布和爆炸延迟时间。结果推进剂慢烤47 h后达到临界温度，其值为352℃；快烤推进剂加热697 s后达到临界温度，临界温度为355℃。结论推进剂在快速烤燃模式下的热扩散速率大于慢速烤燃工况下，但是温度梯度则相反。两种工况下推进剂达到临界温度后开始反应的位置不同，推进剂厚度决定了其储热能力。

The uncertain environmental consequences and regulations associated with using open burning/open detonation for the disposal of energetic materials are forcing both manufacturers and users to examine alternative disposal technologies. In general, these alternatives involve a material removal operation followed by processing steps that lead to reuse of valuable constituents and/or disposal of waste. While a number of post-removal processing options appear to be viable, the initial step of removing an energetic material, such as a solid rocket motorpropellant, from its container remains a significant technological challenge. Large rocket motors containing highly energetic propellant, hazard class 1.1, are of particular concern because of their inherent handling hazards. We will describe the results of a study using thermal cycling to increase the surface area of inert propellant formulations. The propellant removal method studied employs thermal cycling to cryogenic temperatures (cryocycling). Using inert propellants and liquid nitrogen we have demonstrated that this process produces multiple cracks throughout the bulk of the grain. The properties of the actual and inert propellants are being measured, and a model is being developed to relate experiments on inert material to actual propellant. Possible methods to increase thermal gradients, crack propagation and initiation are also presented.

A method of distribution of an additional solid-phase component (oxidizer) providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellantmotor is proposed in the paper.

A method of distribution of an additional solid-phase component (oxidizer) providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellantmotor is proposed in the paper.

The dynamic response properties of the space shuttle solid rocket moter (TP-H1148) propellant were characterized and the expected limits of propellant variability were established. Dynamic shear modulus tests conducted on six production batches of TP-H1148 at various static and dynamic strain levels over the temperature range from 40 F to 90 F. A heat conduction analysis and dynamic response analysis of the space shuttle solid rocket motor (SRM) were also conducted. The dynamic test results show significant dependence on static and dynamic strain levels and considerable batch-to-batch and within-batch variability. However, the results of the SRM dynamic response analyses clearly demonstrate that the stiffness of the propellant has no consequential on the overall SRM dynamic response. Only the mass of the propellant needs to be considered in the dynamic analysis of the space shuttle SRM.

The methods of FEM Modelling and Oscillation Analysis of SolidPropellant Rocket Motor were explores. The FEM model of flexible nozzle using the equivalent model of flexible joint based on the three-direction custom spring elements is build; modify the FEM model of the flexible joint according to the experiment data; and carry out the oscillation analysis of solidpropellant rocket motor.%研究了固态火箭发动机柔性喷管有限元建模及摆动分析方法.利用基于自定义三向弹簧单元的柔性接头线性等效模型,建立了发动机柔性喷管有限元模型.根据试验数据对柔性接头模型进行修正,并对发动机柔性喷管进行了摆动分析.

Pulsed radiography was applied to observe solidpropellant surface regression during rocket motor operation. Using a 150 KV flash X-ray system manufactured by the Field Emission Corporation and two kinds of film suppliers, images of the propellant surface of a 5 cm diameter end burning rocket motor were recorded on film. The repetition frame rate of 8 pulses per second and the pulse train length of 10 pulses are limited by the capability of the power supply and the heat build up within the X-ray tube, respectively. The experiment demonstrated the effectiveness of pulsed radiography for observing solidpropellant surface regression. Measuring the position of burning surface images on film with a microdensitometer, quasi-instantaneous burning rate as a function of pressure and the variation of characteristic velocity with pressure and gas stay time were obtained. Other research items to which pulsed radiography can be applied are also suggested.

The development of new high-performance propellant combinations requires the establishment of safety and handling characteristics and thermodynamic decomposition and explosive properties. This paper addresses the early development phases of a new composite solidpropellant based on HNF as oxidizer a

The calculation of solidpropellant exhaust plume flow fields is addressed. Two major areas covered are: (1) the applicability of empirical data currently available to define particle drag coefficients, heat transfer coefficients, mean particle size and particle size distributions, and (2) thermochemical modeling of the gaseous phase of the flow field. Comparisons of experimentally measured and analytically predicted data are made. The experimental data were obtained for subscale solidpropellantmotors with aluminum loadings of 2, 10 and 15%. Analytical predictions were made using a fully coupled two-phase numerical solution. Data comparisons will be presented for radial distributions at plume axial stations of 5, 12, 16 and 20 diameters.

The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

High burn rate propellants help maintain high levels of thrust without requiring complex, high surface area grain geometries. Utilizing high burn rate propellants allows for simplified grain geometries that not only make production of the grains easier, but the simplified grains tend to have better mechanical strength, which is important in missiles undergoing high-g accelerations. Additionally, high burn rate propellants allow for a higher volumetric loading which reduces the overall missile's size and weight. The purpose of this study is to present methods of achieving a high burn rate propellant and to develop a composite propellant formulation that burns at 1.5 inches per second at 1000 psia. In this study, several means of achieving a high burn rate propellant were presented. In addition, several candidate approaches were evaluated using the Kepner-Tregoe method with hydroxyl terminated polybutadiene (HTPB)-based propellants using burn rate modifiers and dicyclopentadiene (DCPD)-based propellants being selected for further evaluation. Propellants with varying levels of nano-aluminum, nano-iron oxide, FeBTA, and overall solids loading were produced using the HTPB binder and evaluated in order to determine the effect the various ingredients have on the burn rate and to find a formulation that provides the burn rate desired. Experiments were conducted to compare the burn rates of propellants using the binders HTPB and DCPD. The DCPD formulation matched that of the baseline HTPB mix. Finally, GAP-plasticized DCPD gumstock dogbones were attempted to be made for mechanical evaluation. Results from the study show that nano-additives have a substantial effect on propellant burn rate with nano-iron oxide having the largest influence. Of the formulations tested, the highest burn rate was a 84% solids loading mix using nano-aluminum nano-iron oxide, and ammonium perchlorate in a 3:1(20 micron: 200 micron) ratio which achieved a burn rate of 1.2 inches per second at 1000

The results of a one year study to develop a dynamic response model for the Space Shuttle Solid Rocket Motor (SRM) propellant are presented. An extensive literature survey was conducted, from which it was concluded that the only significant variables affecting the dynamic response of the SRM propellant are temperature and frequency. Based on this study, and experimental data on propellants related to the SRM propellant, a dynamic constitutive model was developed in the form of a simple power law with temperature incorporated in the form of a modified power law. A computer program was generated which performs a least-squares curve-fit of laboratory data to determine the model parameters and it calculates dynamic moduli at any desired temperature and frequency. Additional studies investigated dynamic scaling laws and the extent of coupling between the SRM propellant and motor cases. It was found, in agreement with other investigations, that the propellant provides all of the mass and damping characteristics whereas the case provides all of the stiffness.

A test program is described which determines the extent of physical property changes that result from extended space exposure. Primary emphasis was placed on determining the effects of space vacuum. Solidpropellants were stored and their physical properties tested in a vacuum and in a dry environment. The storage caused significantly greater increases in the propellants' modulus and maximum tensile strength than occurred in parallel ambient stored samples. The data indicate that the loss of trace amounts of residual moisture from cured propellant is the apparent cause of the observed propellant property changes. Therefore, initial screening tests were carried out under dry storage conditions. Upon completion of the dry storage tests, appropriate propellant samples are exposed to an actual space environment using the Long Duration Exposure Facility (LDEF).

An overview is presented of experimental methods for determining the combustion-stability properties of solidpropellants. The methods are generally based on either the temporal response to an initial disturbance or on external methods for generating the required oscillations. The size distribution of condensed-phase combustion products are characterized by means of the experimental approaches. The 'T-burner' approach is shown to assist in the derivation of pressure-coupled driving contributions and particle damping in solid-propellant rocket motors. Other techniques examined include the rotating-valve apparatus, the impedance tube, the modulated throat-acoustic damping burner, and the magnetic flowmeter. The paper shows that experimental methods do not exist for measuring the interactions between acoustic velocity oscillations and burning propellant.

Cryogenic solidpropellants (CSP) are a new kind of chemical propellants that use frozen products to ensure the mechanical resistance of the grain. The objective is to combine the high performances of liquid propulsion and the simplicity of solid propulsion. The CSP concept has few disadvantages. Storability is limited by the need of permanent cooling between motor loading and firing. It needs insulations that increase the dry mass. It is possible to limit significantly these drawbacks by using a cooling temperature near the ambient one. It will permit not to change the motor materials and to minimize the supplementary dry mass due to insulator. The designation "Refrigerated SolidPropellant" (RPS) is in that case more appropriate as "Cryogenic SolidPropellant." SNPE Matériaux Energétiques is developing new concept of composition e e with cooling temperature as near the ambient temperature as possible. They are homogeneous and the main ingredients are hydrogen peroxide, polymer and metal or metal hydride, they are called "HydroxalaneTM." This concept allows reaching a high energy level. The expected specific impulse is between 355 and 375 s against 315 s for hydroxyl-terminated polybutadiene (HTPB) / ammonium perchlorate (AP) / Al composition. However, the density is lower than for current propellants, between 1377 and 1462 kg/m3 compared to around 1800 kg/m3 . This is an handicap only for volume-limited application. Works have been carried out at laboratory scale to define the quality of the raw materials and the manufacturing process to realize sample and small grain in a safer manner. To assess the process, a small grain with an internal bore had been realized with a composition based on aluminum and water. This grain had shown very good quality, without any defect, and good bonding properties on the insulator.

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Handling of solidpropellants. 420.65....65 Handling of solidpropellants. (a) A launch site operator shall determine the maximum total quantity of solidpropellants and other solid explosives by class and division, in accordance with 49...

Significant advances have been made during the last decade in several fields of solid propulsion: the advances have enabled new savings in the motor development phase and recurring costs, because they help limit the number of prototypes and tests. The purpose of the paper is to describe the improvements achieved by SNPE in solid grain technologies, making these technologies available for new developments in more efficient and reliable future SRMs: new energetic molecules, new solidpropellants, new processes for grain manufacturing, quick response grain design tools associated with advanced models for grain performance predictions. Using its expertise in chemical synthesis, SNPE develops new molecules to fit new energetic material requirements. Tests based on new propellant formulations have produced good results in the propellant performance/safety behavior ratio. New processes have been developed simultaneously to reduce the manufacturing costs of the new propellants. In addition, the grain design has been optimized by using the latest generation of predictive theoretical tools supported by a large data bank of experimental parameters resulting from over 30 years' experience in solid propulsion: Computer-aided method for the preliminary grain design Advanced models for SRM operating and performance predictions

to use the collision function form proposed by Golovin to simplify this production term: 4C><=) <P- .: Accordingly: m hence, by integration: Now, we...November 21, 1940 in Paris, Seine. VFirst Thesis. "Contribution to the Study of Specific i Impulse Loss in SolidPropellant Rockets." Second Thesis

Solidpropellant is the highly energetic fuel burnt in the combustion chamber of ballistic weapons. It is manufactured, for this purpose, in either granular or stick form. Internal ballistics describes the behavior within the combustion chamber throughout the ballistic cycle upto projectile exit from the muzzle of the gun barrel. Over the last twenty years this has been achieved by modelling the process using two-phase flow equations. The solid granules or sticks constitute ...

Contributions to the evolution of solid-propellant rocketry have come from a variety of sources. World War II research on large solids enabled one company to capitalize on work in the area of castable double-base propellants. Separate development of castable composite propellants led to production of Polaris and Minuteman powerplants. Pivotal to the development of these missiles were Edward Hall's advocacy of the Minuteman missile within the Air Force and contract funding to resolve problems. The discovery that adding large amounts of aluminum significantly increased the specific impulse of a castable composite propellant further aided large-missile technology. These separate lines of research led to the development of large solid-propellantmotors and boosters. Many more discoveries went into the development of large solid-propellantmotors. Ammonium perchlorate replaced potassium perchlorate as an oxidizer in the late 1940's, and binders were developed. Discoveries important in the evolution of large solid-propellantmotors appear to have resulted from innovators' education and skills, an exposure to contemporary problems, an awareness of theory but a willingness not to let it dictate empirical investigations, and proper empirical techniques. Other important contributions are the adequate funding and exchange of information. However, many questions remain about these and other innovations.

China has undertaken to research and develop composite solidpropellant rocket motors since 1958. At the request of the development of space technology, composite solidpropellant rocket motor has developed from small to large, step by step. For the past thirty eight years, much progress has made, many technical obstacles, such as motor design, case materials and their processing technology, propellant formulations and manufacture, nozzles and thrust vector control, safe ignition, environment tests, nondestructive inspection and quality assurance, static firing test and measurement etc. have been solved. A serial of solid rocket motors have been offered for China's satellites launch. The systems of research, design, test and manufacture of solid rocket motors have been formed.

Measurement data on the performance of Space Shuttle Solid Rocket Motor show wide variations in the head-end pressure changes and the total thrust build-up during the ignition transient periods. To analyze the flow and thermal behavior in the tested solid rocket motors, a 1-dimensional, ideal gas flow model via the SIMPLE algorithm was developed. Numerical results showed that burning patterns in the star-shaped head-end segment of the propellant and the erosive burning rate are two important factors controlling the ignition transients. The objective of this study is to extend the model to include the effects of aluminum particle commonly used in solidpropellants. To treat the effects of aluminum-oxide particles in the combustion gas, conservation of mass, momentum, and energy equations for the particles are added in the numerical formulation and integrated by an inter-phase-slip algorithm.

National Aeronautics and Space Administration — ET Materials, LLC developed the first ever electrically controlled extinguishable solidpropellant (ECESP). The original propellant developed under Air Force SBIR...

The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

As the first step of the study of the combustion control of solidpropellants by electrical discharges, the effects of an arc discharge, which flows along the burning surface, on the burning rate and on the increase of enthalpy of the combustion product were investigated. For specially devised composite propellants, which are composed of Al and Teflon powders, it was shown that the combination can be controlled by an arc discharge; the combustion continues when the arc discharge is applied and is interrupted when the arc discharge breaks. In the present investigation, it was also shown that an arc discharge coupled with a high-frequency electrical discharge has potential as an effective ignition method for solidpropellants. For the application of this type of combustion control to an ignitor for a solidpropellant rocket motor or to a control rocket motor, this method lacks flexibility in the configuration scale and needs relatively high electric power at the present stage.

Full Text Available A method of distribution of an additional solid-phase component (oxidizer providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellantmotor is proposed in the paper.

This paper illustrates experiments that were conducted on the formation of solid hydrogen particles in liquid helium. Solid particles of hydrogen were frozen in liquid helium, and observed with a video camera. The solid hydrogen particle sizes, their molecular structure transitions, and their agglomeration times were estimated. article sizes of 1.8 to 4.6 mm (0.07 to 0. 18 in.) were measured. The particle agglomeration times were 0.5 to 11 min, depending on the loading of particles in the dewar. These experiments are the first step toward visually characterizing these particles, and allow designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

A measurement system to study shock initiation behavior of solidpropellants was established experimentally. By using this system, the study on shock initiation to the recovered solidpropellants with micro damage was performed, especially on the deflagration to denonation transition (DDT) process of solidpropellants under both the strong and weak conditions of restriction. The experimental results show that there is a fully compression region in DDT process.

Several experiments on the formation of solid hydrogen particles in liquid helium were recently conducted at the NASA Glenn Research Center at Lewis Field. The solid hydrogen experiments are the first step toward seeing these particles and determining their shape and size. The particles will ultimately store atoms of boron, carbon, or hydrogen, forming an atomic propellant. Atomic propellants will allow rocket vehicles to carry payloads many times heavier than possible with existing rockets or allow them to be much smaller and lighter. Solid hydrogen particles are preferred for storing atoms. Hydrogen is generally an excellent fuel with a low molecular weight. Very low temperature hydrogen particles (T < 4 K) can prevent the atoms from recombining, making it possible for their lifetime to be controlled. Also, particles that are less than 1 mm in diameter are preferred because they can flow easily into a pipe when suspended in liquid helium. The particles and atoms must remain at this low temperature until the fuel is introduced into the engine combustion (or recombination) chamber. Experiments were, therefore, planned to look at the particles and observe their formation and any changes while in liquid helium.

Card gap test, which is standardized in Japan Explosives Society, was modified in order to apply it to solid rocket propellants and carried out to evaluate sensitivities against shock stimuli. Solidpropellants tested here were mainly azide polymer composite propellants, which contained ammonium nitrate (AN) as a main oxidizer. Double base propellant, composed nitroglycerin and nitrocellulose (NC), and ammonium perchlorate (AP)-based composite propellants. It is found that the sensitivity was dominated by the oxidizer characteristics. AP- and AN-based propellant had less sensitivity and HMX-based propellant showed higher sensitivity, and the adding of NC and TMETN contributed to worse sensitive for the card gap test. Good relationship was obtained between the card gap sensitivity and the oxygen balance of propellants tested here. (orig.)

The development status of the solidpropellant engine (P230) of the Ariane-5 launcher is described. Large new industrial plants were built in Europe and Guiana for the development and manufacture of the solid-booster stage and are now operational. A product assurance policy, specific and common to the companies that are involved in the engine's development, was defined and will be implemented. The paper describes the production cycles for the charged segments, the igniter, and the nozzle for P230 engine, as well as the process of engine integration and testing. Consideration is also given to the engine thrust capability, the launcher flight control, and the interfaces. The the major engine development tests are described.

Full Text Available Mechanical properties of solid rocket propellants are dependent on temperature. Any change in temperature brings significant change in the tensile strength, percentage elongation, and elastic modulus of the propellant. Different classes of operational solid rocket propellants namely extruded double-base propellants, composite, extruded composite and nitrarte ester polyester propellants were evaluated at different temperatures in the operating range of the rockets and missiles preferably in the range of –50 oC to +55 oC. It was observed that for each class of propellant, as temperature reduces, propellant becomes hard. This is depicted by increase in elastic modulus and tensile strength of the material. However, trend of percentage elongation is not very uniform. Extruded double-base propellants show less percentage elongation (around 1 per cent at reduced temperature (–50 oC probably due to brittleness. So is the trend with case-bonded composite propellants. However, reverse trend is exhibited by cartridge-loaded composite propellants and nitrate ester polyester propellants. Such propellants show higher percentage elongation (6 per cent for CLCP and 35 per cent for NEPE at reduced temperature (–50 oC. This makes such propellants tough and more area under stress-strain curve at reduced temperature is observed.Defence Science Journal, 2011, 61(6, pp.529-533, DOI:http://dx.doi.org/10.14429/dsj.61.774

Full Text Available A new approach to solve the geometry-problem of solidpropellant star is presented. The basis of the approach is to take the web-thickness (a ballistic as well as a geometrical property as the characteristic length. The nondimensional characteristic parameters representing diameter, length, slenderness-ratio, and ignitor accommodation of the grain are all identified. Many particular cases of star configurations (from the configurations of single propellant to those of four different propellants can be analysed through the identified characteristic parameters. A better way of representing the single-propellant-star-performance in a design graph is explained. Two types of dual propellant grains are analysed in detail. The first type is characterised by its two distinct stages of burning (initially by single propellant burning and then by dual propellant burning; the second type has the dual propellant burning throughout. Suitability of the identified characteristic parameters to an optimisation study is demonstrated through examples.

An analysis of experimental nozzle, exhaust plume, and exhaust plume impingement data is presented. The data were obtained for subscale solidpropellantmotors with propellant Al loadings of 2, 10 and 15% exhausting to simulated altitudes of 50,000, 100,000 and 112,000 ft. Analytical predictions were made using a fully coupled two-phase method of characteristics numerical solution and a technique for defining thermal and pressure environments experienced by bodies immersed in two-phase exhaust plumes.

The recent research on the mechanical properties, burning behavior and processing technology of solidpropellants abroad was reviewed. There are some available results in predicting theoretically the mechanical and rheological properties of solidpropellants. In order to reduce the cost and increase the reliability in propellants processing, there is great demand on the design and manufacture of continuous mixer of high efficiency and safety. The research on the thermoplastic elastomers used as a kind of future binder of solidpropellants has attracted more and more attention of many relevant experts.

This work was conducted to define further the effects of propellant composition variables on the acceleration-induced burning rate augmentation of solidpropellants. The rate augmentation at a given acceleration was found to be a nonlinear inverse function of the reference burning rate and not controlled by binder or catalyst type at a given reference rate. A nonaluminized propellant and a low rate double-base propellant exhibited strong transient rate augmentation due to surface pitting resulting from the retention of hot particles on the propellant surface.

A computerized mathematical model of the combustion response function of composite solidpropellants was developed with particular attention to the contributions of the solid phase heterogeneity. The one-dimensional model treats the solid phase as alternating layers of ammonium perchlorate and binder, with an exothermic melt layer at the surface. Solution of the Fourier heat equation in the solid provides temperature and heat flux distributions with space and time. The problem is solved by conserving the heat flux at the surface from that produced by a suitable model of the gas phase. An approximation of the BDP flame model is utilized to represent the gas phase. By the use of several reasonable assumptions, it is found that a significant portion of the problem can be solved in closed form. A method is presented by which the model can be applied to tetramodal particle size distributions. A computerized steady-state version of the model was completed, which served to validate the various approximations and lay a foundation for the combustion response modeling. The combustion response modeling was completed in a form which does not require an iterative solution, and some preliminary results were acquired.

Full Text Available The ultrasonic pulse-echo technique has been applied for the measurement of instantaneous burnrate of aluminised composite solidpropellants. The tests have been carried out on end-burning 30 mmthick propellant specimens at nearly constant pressure of about 1.9 MPa. Necessary software forpost-test data processing and instantaneous burn rate computations have been developed. The burnrates measured by the ultrasonic technique have been compared with those obtained from ballisticevaluation motor tests on propellant from the same mix. An accuracy of about +- 1 per cent ininstantaneous burn rate measurements and reproducibility of results have been demonstrated byapplying ultrasonic technique.

Physical and chemical ProPerties and combustion characteristics of propellants differ according to the combination of oxidizers and fuel components. Composite smoke propellant, having crystalline ammonium perchlorate as an oxidizer and hydrocarbon Polymer as a fuel, has higher specific impulse and improved mechanical properties compared to smokeless double base propellant consisting of nitroglycerin and nirocellulose. Double base propellants with low specific impulse are combined with nitramines( RDX or HMX ) to make composite modified double based( CMDB ) propellants, as a result the smokeless property of double base propellant is preserved and the combustion efficiency is increased. With the combination of oxidizing agents and fuels, formation of various high functional propellants has been possible and energetic azide polymers have provided possibilities for fuels of propellants. 3 refs., 6 figs., 3 tabs.

The application of laser-induced fluorescence (LIF) to the study of high pressure solidpropellant flames is described. The distribution of the OH and CN radicals was determined in several solidpropellant flames at pressures up to 3.5 MPa. The greatest difficulty in these measurements was the separation of the desired LIF signals from the large scattering at the laser wavelength from the very optically thick propellant flames. Raman experiments using 308-nm excitation were also attempted in the propellant flames but were unsuccessful due to LIF interferences from OH and NH.

Full Text Available Burning rate of a solid rocket propellant depends on pressure and temperature. Conventional strand burner and Crawford bomb test on propellant strands was conducted to assess these dependent parameters. However, behaviour of propellant in rocket motor is different from its behaviour in strand form. To overcome this anomaly, data from static evaluation of rocket motor was directly used for assessment of these burningrate controlling parameters. The conventional empirical power law (r=aoexp[p{T-To}]Pn was considered and a method was evolved for determination of pressure index (n and temperature sensitivity coefficient (p of burning rate for solid rocket propellants from static evaluation data. Effect of pressure index and temperature sensitivity coefficient on firing curve is also depicted. Propellant grain was fired in progressive mode to cover a very wide pressure range of 50 kg/cm2 to 250 kg/cm2 and propellant burning rate index was calculated to be 0.32 in the given pressure range. Propellant grain was fired at +35 °C and 20 °C temperatures and temperature sensitivity coefficient of burning rate was calculated to be 0.27 % per °C. Since both the values were evaluated from realised static evaluation curves, these are more realistic and accurate compared to data generated by conventional methods.Defence Science Journal, 2009, 59(6, pp.666-669, DOI:http://dx.doi.org/10.14429/dsj.59.1573

Pressure coupled response is one of the main causes of combustion instability in the solid rocket motor. It is also a characteristic parameter for predicting the stability. The pressure coupled response function is usually measured by different methods to evaluate the performance of new propellant. Based on T-burner and "burning surface doubled and secondary attenuation", an improved method for measuring the pressure coupled response of composite propellant is introduced in this article. A computational fluid dynamics (CFD) study has also been conducted to validate the method and to understand the pressure oscillation phenomenon in T-burner. Three rounds of tests were carried out on the same batch of aluminized AP/HTPB composite solidpropellant. The experimental results show that the sample propellant had a high response function under the conditions of high pressure (~11.5 MPa) and low frequency (~140 Hz). The numerically predicted oscillation frequency and amplitude are consistent with the experimental results. One practical solid rocket motor using this sample propellant was found to experience pressure oscillation at the end of burning. This confirms that the sample propellant is prone to combustion instability. Finally, acoustic pressure distribution and phase difference in T-burner were analyzed. Both the experimental and numerical results are found to be associated with similar acoustic pressure distribution. And the phase difference analysis showed that the pressure oscillations at the head end of the T-burner are 180° out of phase from those in the aft end of the T-burner.

The crew launch vehicle considered for the Constellation (Cx) Program utilizes a first stage solid rocket motor. If an abort is initiated in first stage flight the Crew Module (CM) will separate and be pulled away from the launch vehicle via a Launch Abort System (LAS) in order to safely and quickly carry the crew away from the malfunction launch vehicle. Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solidpropellant, poses a threat to the CM. The harsh radiative thermal environment induced by surrounding burning propellant debris may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and quantify the risk of first stage propellant debris leading to radiative thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst-case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature, with magnitudes on the order of 10 s of degrees Fahrenheit. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be

The goal of the SolidPropellant Resupply Team is to develop Crusader system concepts capable of automatically handling 155mm projectiles and Modular Artillery Charges (MACs) based on system requirements. The system encompasses all aspects of handling from initial input into a resupply vehicle (RSV) to the final loading into the breech of the self-propelled howitzer (SPH). The team, comprised of persons from military and other government organizations, developed concepts for the overall vehicles as well as their interior handling components. An intermediate review was conducted on those components, and revised concepts were completed in May 1995. A concept evaluation was conducted on the finalized concepts, from both a systems level and a component level. The team`s Best Technical Approach (BTA) concept was selected from that evaluation. Both vehicles in the BTA have a front-engine configuration with the crew situated behind the engine-low in the vehicles. The SPH concept utilizes an automated reload port at the rear of the vehicle, centered high. The RSV transfer boom will dock with this port to allow automated ammunition transfer. The SPH rearm system utilizes fully redundant dual loaders. Active magazines are used for both projectiles and MACs. The SPH also uses a nonconventional tilted ring turret configuration to maximize the available interior volume in the vehicle. This configuration can be rearmed at any elevation angle but only at 0{degree} azimuth. The RSV configuration is similar to that of the SPH. The RSV utilizes passive storage racks with a pick-and-place manipulator for handling the projectiles and active magazines for the MACs. A telescoping transfer boom extends out the front of the vehicle over the crew and engine.

The goal of the SolidPropellant Resupply Team is to develop Crusader system concepts capable of automatically handling 155mm projectiles and Modular Artillery Charges (MACs) based on system requirements. The system encompasses all aspects of handling from initial input into a resupply vehicle (RSV) to the final loading into the breech of the self-propelled howitzer (SPH). The team, comprised of persons from military and other government organizations, developed concepts for the overall vehicles as well as their interior handling components. An intermediate review was conducted on those components, and revised concepts were completed in May 1995. A concept evaluation was conducted on the finalized concepts, from both a systems level and a component level. The team`s Best Technical Approach (BTA) concept was selected from that evaluation. Both vehicles in the BTA have a front-engine configuration with the crew situated behind the engine-low in the vehicles. The SPH concept utilizes an automated reload port at the rear of the vehicle, centered high. The RSV transfer boom will dock with this port to allow automated ammunition transfer. The SPH rearm system utilizes fully redundant dual loaders. Active magazines are used for both projectiles and MACs. The SPH also uses a nonconventional tilted ring turret configuration to maximize the available interior volume in the vehicle. This configuration can be rearmed at any elevation angle but only at 0{degree} azimuth. The RSV configuration is similar to that of the SPH. The RSV utilizes passive storage racks with a pick-and-place manipulator for handling the projectiles and active magazines for the MACs. A telescoping transfer boom extends out the front of the vehicle over the crew and engine.

Full Text Available Composite solidpropellant is prepared using tri-modal Ammonium perchlorate (AP containing coarse, fine and ultrafine fractions of AP with average particle size (APS 340, 40 and 5 μm respectively, in various compositions and their rheological, mechanical and burn rate characteristics are evaluated. The optimum combination of AP coarse to fine to ultrafine weight fraction was obtained by testing of series of propellant samples by varying the AP fractions at fixed solid loading. The concentration of aluminium was maintained constant throughout the experiments for ballistics requirement. The propellant formulation prepared using AP with coarse to fine to ultrafine ratio of 67:24:9 has lowest viscosity for the propellant paste and highest tensile strength due to dense packing as supported by the literature. A minimum modulus value was also observed at 9 wt. % of ultrafine AP concentration indicates the maximum solids packing density at this ratio of AP fractions. The burn rate is evaluated at different pressures to obtain pressure exponent. Incorporation of ultrafine fraction of AP in propellant increased burn rate without adversely affecting the pressure exponent. Higher solid loading propellants are prepared by increased AP concentration from 67 to 71 wt. % using AP with coarse to fine to ultrafine ratio of 67:24:9. Higher solid content up to 89 wt. % was achieved and hence increased solidmotor performance. The unloading viscosity showed a trend with increased AP content and the propellant couldn't able to cast beyond 71 wt. % of AP. Mechanical properties were also studied and from the experiments noticed that % elongation decreased with increased AP content from 67 to 71 wt.%, whereas tensile strength and modulus increased. Burn rate increased with increased AP content and observed that pressure exponent also increased and it is high for the propellant containing with 71 wt.% of AP due to increased oxidiser to fuel ratio. Catalysed

Special aluminum and ammonium perchlorate (AP) particle size distributions were prepared for a matrix of five-inch diameter, center-perforated (CP) motor tests to measure the aluminum oxide slag response in Space Shuttle Reusable Solid Rocket Motor (RSRM) propellant. Previous tests of TP-H1148 propellant in five-inch CP spin motors have shown a correlation between aluminum particle size and generated slag. The motors for this study were cast from thirteen five-gallon propellant mixes which used five particle size levels of aluminum powder, five of unground AP and three of ground AP. Aluminum had the greatest effect on slag formation, the more coarse fractions causing greater slag quantities and larger slag particles. Unground AP had about half the effect of aluminum with the coarser fractions again producing more and larger sized slag particles. The variation in ground AP did not have a significant effect on slag formation. Quench bomb tests showed the same trends as the spin motors, that is, larger aluminum and AP particle size distributions generated larger slag particles leaving the propellant surface. Cured propellant mechanical properties were also impacted by particle size variation.

Full Text Available A generalised model of burning of a solid rocket propellant based on kinetics of propellant hasbeen developed. A complete set of variables has been formed after examining the existing models.Buckingham theorem provides the functional form of the model, such that the existing models are thesubcases of this generalised model. This proposed model has been validated by an experimental data.

The manufacturing sequence of NASA's new Advanced Solid Rocket Motor, developed as a replacement of the Space Shuttle's existing Redesigned Solid Rocket Motor, is overviewed. Special attention is given to the case preparation, the propellant mix/cast, the nondestructuve evaluation, the motor finishing, and the refurbishment. The fabrication sequences of the case, the nozzle, and the igniter are described.

Full Text Available High frequency combustion instabilities imply a major risk for the solid rocket motor stableworking and they are directly linked to the propellant response to chamber pressure coupling. Thisarticle discusses a laboratory testing method for the measurement and evaluation of the pressurecoupled response for non-metalized propellants in a first stage. Experimental researches were donewith an adequate setup, built and improved in our lab, able to evaluate the propellant response byinterpreting the pressure oscillations damping in terms of propellant response. Our paper aims atdefining a linearized one-dimensional flow study model to analyze the disturbed operation of the solidpropellant rocket motors. Based on the applied model we can assert that the real part of propellantresponse is a function of the oscillations damping, acoustic energy in the motor chamber and variouslosses in the burning chamber. The imaginary part of propellant response mainly depends on thenormalized pulsation, on the burning chamber gas column and on the pressure oscillations frequency.Our research purpose was obviously to minimize the risk of the combustion instabilities effects on therocket motors working, by experimental investigations using jet modulating techniques and sustainedby an interesting study model based on the perturbation method.

Through the history of composite solidpropellant binders new chemicals are introduced as binders to improve upon the previous generation. Sometimes this is done to improve upon the flaws or shortcomings of a previous binder. Other time it is to meet a new set of requirements desired by industry. Dicyclopentadiene (DCPD) is a hydrocarbon monomer being considered for its potential as a new binder in the composite propellant industry. The binder of a composite solidpropellant is arguably the most important feature of the propellant. It is the binder that provides the majority of the structural characteristics of the propellant while also contributing itself as fuel to the combustion process. A binder in composite propellants must also be able to accept the introduction of a large quantity of solid filler; oxidizer, fuel, and other energetic and non-energetic particles. Many of the composite propellants used in industry today have over 80% of their weight composed of non-binder solid or liquid fillers. These requirements must be met by the binder in some form or fashion to produce a propellant able to compete with binders currently in use. When DCPD is polymerized it produces an extremely tough plastic with excellent tensile and impact strength. Experimentation has found that DCPD is able to support a large quantity of solid materials, over 80% weight of the mixture, while still retaining a great portion of its original strength. When compared to another similarly loaded binder currently used in industry, Hydroxyl-Terminated Polybutadiene (HTPB), it was found that DCPD composite propellant had nearly 1.5 times the stress capacity while still exhibiting over 75% of the strain capacity of HTPB based composite propellant. In addition it was also shown that DCPD composite propellant allows for tailoring of its mechanical properties with the addition of plasticizers. The DCPD based composite propellant also exhibits a burning rate nearly twice that HTPB. These factors

This colloquium has been jointly organized by the research center of history of sciences and techniques (CRHST) and the association of the friends of the gunpowder and pyrotechnical patrimony (A3P). It gathers historians of sciences and techniques and specialists of solidpropellants and their applications who make a review of the approaches that have led to todays propellants efficiency and mastery. This books contains 2 introductive talks, 24 articles, a round table and some concluding remarks. The articles deal with: 1 - from the black powder rockets to the space shuttle: France, pioneer of solid propulsion, from Vaillant to Damblanc (1821-1938); the development of solidpropellants in the 20. century; lessons learnt from the Challenger accident; 2 - the institutions: the laboratory of ballistics of Sevran-Livry (1945-1969); an historical overview of ONERA's researches on solidpropellants; the cast propellants at the Direction of Explosives (1945-1955); 3 - the propellants: the manufacturing secrets of the extruded double base propellants; the development of cast double base propellants; the invention of composite propellants; 4 - space applications: the Diamant adventure; the solidpropellant engines of Ariane 5, an endless story; P80, a new generation of solidpropellant engines for space applications; 6 - physics and models: from ap{sup n} to 3-D simulations: the combustion of solidpropellants in the 20. century; the mechanical behaviour of solidpropellant loads (1960-70 years); composite propellants and static electricity (SE) or the occurrence of SE in the manufacturing and implementation of composite propellants; a priori calculation of the performances and synthesis of new energy materials for propellants; 6 - defense applications: French solidpropellant rockets and missiles up to the 1960's; from PHI 1500 to PHI 1930 or the fabulous history of metallic and roving propulsion systems; the G2P, the propulsion system of the M4, the exploratory

area of the propellant grain satisfies the designed value. But cracks in propellant grain can be generated during manufacture, storage, handing and so on. The cracks can provide additional surface area for combustion. The additional combustion may significantly deviate the performance of the rocket motor from the designed conditions, even lead to explosive catastrophe. Therefore a thorough study on the combustion, propagation and fracture of solidpropellant cracks must be conducted. This paper takes an isolated propellant crack as the object and studies the effect of chamber pressurization rate on the combustion, propagation and fracture of the crack by experiment and theoretical calculation. deformable, the burning inside a solidpropellant crack is a coupling of solid mechanics and combustion dynamics. In this paper, a theoretical model describing the combustion, propagation and fracture of the crack was formulated and solved numerically. The interaction of structural deformation and combustion process was included in the theoretical model. The conservation equations for compressible fluid flow, the equation of state for perfect gas, the heat conducting equation for the solid-phase, constitutive equation for propellant, J-integral fracture criterion and so on are used in the model. The convective burning inside the crack and the propagation and fracture of the crack were numerically studied by solving the set of nonlinear, inhomogeneous gas-phase governing equations and solid-phase equations. On the other hand, the combustion experiments for propellant specimens with a precut crack were conducted by RTR system. Predicted results are in good agreement with experimental data, which validates the reasonableness of the theoretical model. Both theoretical and experimental results indicate that the chamber pressurization rate has strong effects on the convective burning in the crack, crack fracture initiation and fracture pattern.

In this paper, described is the development of a numerical simulation system, what we call “Advanced Computer Science on SRM Internal Ballistics (ACSSIB)”, for the purpose of improvement of performance and reliability of solid rocket motors (SRM). The ACSSIB system is consisting of a casting simulation code of solidpropellant slurry, correlation database of local burning-rate of cured propellant in terms of local slurry flow characteristics, and a numerical code for the internal ballistics of SRM, as well as relevant hardware. This paper describes mainly the objectives, the contents of this R&D, and the output of the fiscal year of 2008.

Full Text Available Aluminum-water reactions have been proposed and studied for several decades for underwater propulsion systems and applications requiring hydrogen generation. Aluminum and water have also been proposed as a frozen propellant, and there have been proposals for other refrigerated propellants that could be mixed, frozen in situ, and used as solidpropellants. However, little work has been done to determine the feasibility of these concepts. With the recent availability of nanoscale aluminum, a simple binary formulation with water is now feasible. Nanosized aluminum has a lower ignition temperature than micron-sized aluminum particles, partly due to its high surface area, and burning times are much faster than micron aluminum. Frozen nanoscale aluminum and water mixtures are stable, as well as insensitive to electrostatic discharge, impact, and shock. Here we report a study of the feasibility of an nAl-ice propellant in small-scale rocket experiments. The focus here is not to develop an optimized propellant; however improved formulations are possible. Several static motor experiments have been conducted, including using a flight-weight casing. The flight weight casing was used in the first sounding rocket test of an aluminum-ice propellant, establishing a proof of concept for simple propellant mixtures making use of nanoscale particles.

Rocket test firings were performed to measure the transition length threshold conditions while systematically varying various rocket motor parameters. These include the crossflow velocity, the chamber pressure, the propellant nonerosive burning rate, the propellant surface roughness, and the motor port diameter. The erosive burning trends with varying propellant burning rate, motor chamber pressure, and mass flow rate are consistent with published results.

The purpose of this paper is to characterize the regression rate behavior of hybrid rocket motorpropellant combinations, using hydrogen peroxide (HP), gaseous oxygen (GOX), nitrous oxide (N2O) as the oxidizer and hydroxyl-terminated poly-butadiene (HTPB) as the based fuel. In order to complete this research by experiment and simulation, a hybrid rocket motor test system and a numerical simulation model are established. Series of hybrid rocket motor firing tests are conducted burning different propellant combinations, and several of those are used as references for numerical simulations. The numerical simulation model is developed by combining the Navies-Stokes equations with the turbulence model, one-step global reaction model, and solid-gas coupling model. The distribution of regression rate along the axis is determined by applying simulation mode to predict the combustion process and heat transfer inside the hybrid rocket motor. The time-space averaged regression rate has a good agreement between the numerical value and experimental data. The results indicate that the N2O/HTPB and GOX/HTPB propellant combinations have a higher regression rate, since the enhancement effect of latter is significant due to its higher flame temperature. Furthermore, the containing of aluminum (Al) and/or ammonium perchlorate(AP) in the grain does enhance the regression rate, mainly due to the more energy released inside the chamber and heat feedback to the grain surface by the aluminum combustion.

Full Text Available Due to the continuous run for a green environment the current article proposes a new type of solidpropellant based on the fairly new synthesized oxidizer, ammonium dinitramide (ADN. Apart of having a higher specific impulse than the worldwide renowned oxidizer, ammonium perchlorate, ADN has the advantage, of leaving behind only nitrogen, oxygen and water after decomposing at high temperatures and therefore totally avoiding the formation of hydrogen chloride fumes. Based on the oxidizer to fuel ratios of the current formulations of the major rocket solid booster (e.g. Space Shuttle’s SRB, Ariane 5’s SRB which comprises mass variations of ammonium perchlorate oxidizer (70-75%, atomized aluminum powder (10-18% and polybutadiene binder (12-20% a new solidpropellant was formulated. As previously stated, the new propellant formula and its variations use ADN as oxidizer and erythritol tetranitrate as fuel, keeping the same polybutadiene as binder.

Full Text Available The study reports the effect of incorporation of Al and ammonium perchlorate (AP individually and in combination with each other on combustion pattern and specific impulse (Isp of minimum signature propellants. Incorporation of Al obviates the combustion instability problems; however, it has marginal effect on burning rates. The composition containing AP and zirconium silicate combination gives superior performance; however, its Isp is considerably lower than the composition incorporating 9 per cent AP. A combination of 6 per cent Al gave 20 per cent enhancement in burning rate and 12 s increase in Isp as compared to purely nitramine-based composition, cal-val results also reveal increase in energy output on incorporating AP and Al. Hot stage microscopic and propellant combustion studies indicate occurrence of intense decomposition reaction in case of AP-based compositions.

Full Text Available The influence of constant and transient radiant flux on the burning rate of solidpropellants is considered. The validity of the equivalence principle for the radiant flux and increase in initial temperature and also the problem of possible photochemical effect of thermal radiation are discussed. Experimental data on burning rate response to periodical perturbations of radiant flux for different types of solidpropellants are reported. The problem of correlation between burning rate response to perturbations of pressure and external radiation is considered. Formulation of the problem on transient combustion in terms of the Zeldovich- Novozhilov phenomenological approach is described and the results of numerical integration are presented.

Full Text Available Heat-up times derived from studies on the ignition characteristics of a few model composite solidpropellants, containing polystyrene, carboxy-terminated polybutadiene, plasticised polyvinyl chloride and polyphenol formaldehyde as binders, show that they are directly proportional to the mass of the sample and inversely proportional to the heat flux. Propellant weight-loss prior to ignition and high pressure ignition temperature data on the propellants, ammonium per chlorate, and binders show that the ignition is governed by the gasification of the binder pyrolysis products. The activation energy for the gasification of the pyrolysed polymer products corresponds to their ignition behaviour suggesting that propellant ignition is controlled by the binder.

During the launch of STS-54, a 15 psi pressure blip was observed in the ballistic pressure trace of one of the two Space Shuttle Redesigned Solid Rocket Motors (RSRM). One possible scenario for the observed pressure increase deals with aluminum oxide slag formation in the RSRM. The purpose of this investigation was to examine changes which may have occurred in the aluminum oxide formation in shuttle solidpropellant due to changes in the ammonium perchlorate. Aluminum oxide formation from three propellants, all having the same formulation, but containing ammonium perchlorate from different manufacturers, will be compared. Three methods have been used to look for possible differences among the propellants. The first method was to examine window bomb movies of the propellants burning at 100, 300 and 600 psia. The motor operating pressure during the pressure blip was around 600 psia. The second method used small samples of propellant which were fired in a combustion bomb which quenched the burning aluminum particles soon after they left the propellant surface. The bomb was fired in both argon and Nitrogen atmospheres at various pressures. Products from this device were examined by optical microscopy. The third method used larger propellant samples fired into a particle collection device which allowed the aluminum to react and combust more completely. This device was pressurized with Nitrogen to motor operating pressures. The collected products were subdivided into size fractions by screening and sedimentation and analyzed optically with an optical microscope. the results from all three methods indicate very small changes in the size distribution of combustion products.

4~ .A*4 ~.Zwe SOUMVV Ch.&4 0IVC&TIN 0 e*9 066so. 4 evt’ o R..e High speed, high resolution motion pictures were taken to compare the cinematic data...propellant. High speed, high resolution motion pictures were taken to compare the cinematic data with that available from the holograms. TABLE OF...ally employ finely powered aluminum (1-50 microns) in an attempt to capitalize on the conversion of its high heat of formation to kinetic energy for

The feasibility of a system capable of rapidly and directly measuring the low-frequency (motor characteristics length bulk mode) combustion response characteristics of solidpropellants has been investigated. The system consists of a variable frequency oscillatory driver device coupled with an improved version of the JPL microwave propellant regression rate measurement system. The ratio of the normalized regression rate and pressure amplitudes and their relative phase are measured as a function of varying pressure level and frequency. Test results with a well-characterized PBAN-AP propellant formulation were found to compare favorably with the results of more conventional stability measurement techniques.

Worldwide developments are ongoing to develop new and more energetic composite solidpropellant formulations for space transportation and military applications. Since the 90's, the use of HNF as a new high performance oxidiser is being reinvestigated. Within European development programmes, signific

An Eulerian-Lagrangian two-phase approach was adopted to model the multi-phase reacting internal flow in a solid rocket with a metalized propellant. An Eulerian description was used to analyze the motion of the continuous phase which includes the gas as well as the small (micron-sized) particulates, while a Lagrangian description is used for the analysis of the discrete phase which consists of the larger particulates in the motor chamber. The particulates consist of Al and Al2O3 such that the particulate composition is 100 percent Al at injection from the propellant surface with Al2O3 fraction increasing due to combustion along the particle trajectory. An empirical model is used to compute the combustion rate for agglomerates while the continuous phase chemistry is treated using chemical equilibrium. The computer code was used to simulate the reacting flow in a solid rocket motor with an AP/HTPB/Al propellant. The computed results show the existence of an extended combustion zone in the chamber rather than a thin reaction region. The presence of the extended combustion zone results in the chamber flow field and chemical being far from isothermal (as would be predicted by a surface combustion assumption). The temperature in the chamber increases from about 2600 K at the propellant surface to about 3350 K in the core. Similarly the chemical composition and the density of the propellant gas also show spatially non-uniform distribution in the chamber. The analysis developed under the present effort provides a more sophisticated tool for solid rocket internal flow predictions than is presently available, and can be useful in studying apparent anomalies and improving the simple correlations currently in use. The code can be used in the analysis of combustion efficiency, thermal load in the internal insulation, plume radiation, etc.

Full Text Available Carbon fiber reinforced carbon and silicon carbide matrix composites for nozzle inner of solidpropellant ramjet were prepared by using the hybrid process of "chemical vapor infiltration + precursor impregnation pyrolysis (CVI+PIP". The microstructure, flexural and anti-ablation properties of the C/C-SiC composites and hydraulic test and rocket motor hot firing test for nozzle inner of solidpropellant ramjet were comprehensively investigated. The results show that when the flexural strength of the composite reachs 197 MPa, the fracture damage behavior of the composites presents typical toughness mode.Also the composites has excellent anti-ablative property, i.e., linear ablation rate is only 0.0063 mm·s-1 after 200 s ablation. The C/C-SiC component have excellent integral bearing performance with the hydraulic bursting pressure of 6.5 MPa, and the high temperature combination property of the C/C-SiC composite nozzle inner is verified through motor hot firing of solidpropellant ramjet.

Biocatalytic propulsion is expected to play an important role in the future of micromotors as it might drastically increase the number of available fuelling reactions. However, most of the enzyme-propelled micromotors so far reported still rely on the degradation of peroxide by catalase, in spite of being vulnerable to relatively high peroxide concentrations. To overcome this limitation, herein we present a strategy to encapsulate the catalase and to graft the resulting enzyme capsules on motor particles. Significant improvement of the stability in the presence of peroxide and other aggressive agents has been observed.Biocatalytic propulsion is expected to play an important role in the future of micromotors as it might drastically increase the number of available fuelling reactions. However, most of the enzyme-propelled micromotors so far reported still rely on the degradation of peroxide by catalase, in spite of being vulnerable to relatively high peroxide concentrations. To overcome this limitation, herein we present a strategy to encapsulate the catalase and to graft the resulting enzyme capsules on motor particles. Significant improvement of the stability in the presence of peroxide and other aggressive agents has been observed. Electronic supplementary information (ESI) available. See DOI: 10.1039/c4nr02459a

The depletion rates of NMA and 2-NDPA were investigated by accelerated aging test in NEPE solidpropellant containing BTTN and DEGDN as nitrate ester plasticizers. It was found that both NMA and 2-NDPA were depleted by the 0{sup th} order reactions and these reactions had two kinds of activation energy which showed the break-point at 60 {sup o} C for NMA and 70 {sup o} C for 2-NDPA from Arrhenius plots in temperature range of 20-80 {sup o} C. Therefore, the stabilizer content in propellant could be predicted much better by using reaction rate in the low temperature range than that in the high temperature range. The gas fissuring was not occurred under mild conditions like slow evacuation of gases from decomposition of nitrate ester plasticizers even though NMA and 2-NDPA were completely depleted in propellant.

Double base and composite propellants are generally used for rocket motors, whereby double base propellants basically consist of nitrocellulose plasticized with an explosive plasticizer, mostly nitroglycerine, and in some cases with an additional inert plasticizer and ballistic additives. Composite propellants consist of an oxidizer like ammonium perchlorate and of aluminum, binder and plasticizer and often contain liquid or solid burning rate catalysts. A common feature of both propellants is that they contain smaller or larger amounts of chemically unbonded liquid species which tend to migrate. If these propellants loose part of the plasticizer by migration into the insulation layer, not only will there be a change in mechanical propellant properties but also the bond between propellant and insulation may degrade. However, depending on the severity of these effects, the change in the ballistic properties of the propellant grain caused by plasticizer migration may be of even more importance. In the past, most emphasis was placed on the behaviour of end-burning configurations. However, more recent theoretical and experimental studies revealed that not only for end-burning grain configurations but also for internal burning configurations there is a common effect which is responsible for ballistic anomalies: migration of liquid species from the propellant into the insulation. By using a plasticizer equilibrated insulation in an internal burning configuration the liquid species migration and thus the previously observed ballistic anomalies are avoided. Using this approach for end-burning configurations provides similar positive results. The various factors affecting plasticizer migration are studied and discussed, and several methods to prevent liquid species migration are described as well as methods to obtain plasticizer resistant insulations.

Full Text Available The cure kinetics of propellant slurry based on hydroxy-terminated polybutadiene (HTPB and toluene diisocyanate (TDI polyurethane reaction has been studied by viscosity build up method. The viscosity (ɳ–time (t plots conform to the exponential function ɳ = aebt, where a & b are empirical constants. The rate constants (k for viscosity build up at various shear rate (rpm, evaluated from the slope of dɳ/dt versus ɳ plots at different temperatures, were found to vary from 0.0032 to 0.0052 min-1. It was observed that the increasing shear rate did not have significant effect on the reaction rate constants for viscosity build up of the propellant slurry. The activation energy (Eɳ, calculated from the Arrhenius plots, was found to be 13.17±1.78 kJ mole-1, whereas the activation enthalpy (∆Hɳ* and entropy (∆Sɳ* of the propellant slurry, calculated from Eyring relationship, were found to be 10.48±1.78 kJ mole-1 and –258.51± 5.38 J mole-1K-1, respectively. The reaction quenching temperature of the propellant slurry was found to be -9 ° C, based upon the experimental data. This opens up an avenue for a “freeze-and-store”, then “warm-up and cast”, mode of manufacturing of very large solid rocket propellant grains.

Full Text Available Maxwell fluid model consisting of a spring and a dashpot in series is applied for viscoelastic characterisation of solid rocket propellants. Suitable values of spring constant and damping coefficient wereemployed by least square variation of errors for generation of complete stress-strain curve in uniaxial tensile mode for case-bonded solidpropellant formulations. Propellants from the same lot were tested at different strain rates. It was observed that change in spring constant, representing elastic part was very small with strain rate but damping constant varies significantly with variation in strain rate. For a typical propellant formulation, when strain rate was raised from 0.00037/s to 0.185/s, spring constant K changed from 5.5 MPato 7.9 MPa, but damping coefficient D was reduced from 1400 MPa-s to 4 MPa-s. For all strain rates, stress-strain curve was generated using Maxwell model and close matching with actual test curve was observed.This indicates validity of Maxwell fluid model for uniaxial tensile testing curves of case-bonded solidpropellant formulations. It was established that at higher strain rate, damping coefficient becomes negligible as compared to spring constant. It was also observed that variation of spring constant is logarithmic with strain rate and that of damping coefficient follows power law. The correlation coefficients were introduced to ascertain spring constants and damping coefficients at any strain rate from that at a reference strain rate. Correlationfor spring constant needs a coefficient H, which is function of propellant formulation alone and not of test conditions and the equation developeds K2 = K1 + H ´ ln{(de2/dt/(de1/dt}. Similarly for damping coefficient D also another constant S is introduced and prediction formula is given by D2 = D1 ´ {(de2/dt/(de1/dt}S.Evaluating constants H and S at different strain rates validate this mathematical formulation for differentpropellant formulations

The prototype of a solidpropellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solidpropellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

The combustion temperature of solidpropellant was measured in this paper. Emission spectra of the combustion flame were collected with remote sensing FTIR at the resolution of 4 cm(-1). The combustion temperatures with the burning time were calculated from the maximum spectral line intensity and the molecular rotation-vibration spectra of HF molecule, respectively. Combustion temperatures at each time were all 1 788.8 K from the maximum spectral line intensity method. For comparison, the temperatures calculated from the molecular rotation-vibration spectra were 1 859.7, 1 848. 3, 1 804.0 and 1 782.7 K, respectively. Results show that the two methods are all dependable in measuring combustion temperature of solidpropellant. But the maximum spectral line intensity method is more convenient and rapid than the other when the combustion is relatively stable.

Actin-based motility of Listeria monocytogenes propelled by filament end-tracking molecular motors has been simulated. Such systems may act as potential nanoscale actuators and shuttles useful in sorting and sensing biomolecules. Filaments are modeled as three-dimensional elastic springs distributed on one end of the capsule and persistently attached to the motile bacterial surface through an end-tracking motor complex. Filament distribution is random, and monomer concentration decreases linearly as a function of position on the bacterial surface. Filament growth rate increases with monomer concentration but decreases with the extent of compression. The growing filaments exert push-pull forces on the bacterial surface. In addition to forces, torques arise due to two factors—distribution of motors on the bacterial surface, and coupling of torsion upon growth due to the right-handed helicity of F-actin—causing the motile object to undergo simultaneous translation and rotation. The trajectory of the bacterium is simulated by performing a force and torque balance on the bacterium. All simulations use a fixed value of torsion. Simulations show strong alignment of the filaments and the long axis of the bacterium along the direction of motion. In the absence of torsion, the bacterial surface essentially moves along the direction of the long axis. When a small amount of the torsion is applied to the bacterial surface, the bacterium is seen to move in right-handed helical trajectories, consistent with experimental observations.

Full Text Available A new family of polymeric binders for solid composite propellants is proposed, based on two component interpenetrating polymer networks (IPNs. These networks comprise two different polyurethanes based on hydroxy terminated polybutadiene and ISRO polyol interpenetrated with two different vinyl polymers, viz poly methyl methacrylate and polystyrene. the networks synthesized by the simultaneous interpenetrating technique have been characterized for their properties, such as stress-strain, density, viscosity, thermal degradation, and heat of combustion. Phase morphologies have been determined using electron microscopy. Suitable explanations have been adduced to rationalize the properties of IPNs in terms of their structures and chain interactions. A study of the mechanical properties and burning rates of the ammonium perchlorate (AP-based solidpropellant using the newly synthesised IPNs as binders, has been carried out. The results show that both mechanical strength and burning rate of solidpropellants could be suitably modified by simply changing the nature and/or the ratio of the two interpenetrating polymer components.

In 2002 and 2003, Bradford Engineering B.V. conducted, in corporation with the Dutch research institute TNO Prins Maurits Laboratory (PML) a SME study for ESA-ESTEC for the identification of spaceflight applications and on-ground demonstration of SolidPropellant Cool Gas Generator (SPCGG) technology. This innovative technology has been developed by TNO-PML while Bradford Engineering also brought in its experience in spaceflight hardware development and manufacturing. The SolidPropellant Cool Gas Generator (SPCGG) technology allows for pure gas generation at ambient temperatures, as opposed to conventional solidpropellant gas generators. This makes the SPCGG technology interesting for a wide range of terrestrial spaceflight applications. During the first part of the study, a variety of potential applications have been identified and three applications were selected for a more detailed quantitative study. In the third phase a ground demonstration was performed successfully for a cold gas propulsion system application. During the actual demonstration test, 10 cool gas generators were mounted and all operated successfully in sequence, demonstrating good repeatability of the produced amount of gas and pressure.

Full Text Available This paper presents the procedure of uniaxial mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene (HTPB, whose mechanical properties strongly depend on temperature, strain rate, natural aging and accumulated damage. A method of processing data is presented in order to determine time-temperature shift factor and master curves for tensile strength, ultimate strain and relaxation modulus, depending on reduced time. Functional dependences of these features represent an input for structural analysis of a rocket motorpropellant grain. The effects of natural aging on the mechanical properties are also considered. [Projekat Ministarstva nauke Republike Srbije, br. TR 36050: Research and development of unmanned aircraft in support of traffic infrastructure monitoring

The University of Arizona program is aimed at introducing scientific rigor to the predictability and quality assurance of composite solidpropellants. Two separate approaches are followed: to use the modern analytical techniques to experimentally study carefully controlled propellant batches to discern trends in mixing, casting, and cure; and to examine a vast bank of data, that has fairly detailed information on the ingredients, processing, and rocket firing results. The experimental and analytical work is described briefly. The principle findings were that: (1) pre- (dry) blending of the coarse and fine ammonium perchlorate can significantly improve the uniformity of mixing; (2) the Fourier transformed IR spectra of the uncured and cured polymer have valuable data on the state of the fuel; (3) there are considerable non-uniformities in the propellant slurry composition near the solid surfaces (blades, walls) compared to the bulk slurry; and (4) in situ measurements of slurry viscosity continuously during mixing can give a good indication of the state of the slurry. Several important observations in the study of the data bank are discussed.

An analysis has been performed for oscillatory burning of solidpropellants including gas phase time lag. The gaseous flame is assumed to be premixed and laminar with a one-step overall chemical reaction. The propellant is assumed to decompose according to the Arrenhius Law, with no condensed phase reaction. With this model, strong gas phase resonance has been found in certain cases at the characteristic gas-phase frequencies, but the peaking of the acoustic admittance is in the direction favoring the damping of pressure waves. At still higher frequencies, moderate wave-amplifying ability was found. The limit of low frequency response obtained previously by Denison and Baum was recovered, and the limitations of the quasi-steady theory were investigated.

An analysis has been performed for oscillatory burning of solidpropellants including gas phase time lag. The gaseous flame is assumed to be premixed and laminar with a one-step overall chemical reaction. The propellant is assumed to decompose according to the Arrenhius Law, with no condensed phase reaction. With this model, strong gas phase resonance has been found in certain cases at the characteristic gas-phase frequencies, but the peaking of the acoustic admittance is in the direction favoring the damping of pressure waves. At still higher frequencies, moderate wave-amplifying ability was found. The limit of low frequency response obtained previously by Denison and Baum was recovered, and the limitations of the quasi-steady theory were investigated.

Experiments were conducted to gain insight into the unsatisfactory performance of the salt quench system of solidpropellants in earlier studies. Nine open-air salt spray tests were conducted and high-speed cinematographic coverage was obtained of the events. It is shown that the salt spray by the detonator is generally a two-step process yielding two different fractions. The first fraction consists of finely powdered salt and moves practically unidirectionally at a high velocity (thousand of feet per second) while the second fraction consists of coarse particles and moves randomly at a low velocity (a few feet per second). Further investigation is required to verify the speculation that a lower quench charge ratio (weight of salt/propellant burning area) than previously employed may lead to an efficient quench

In the early 1970's a program was initiated at the Naval Surface Warfare Center/Indian Head Division (NSWC/IHDIV) to address the well-known problems associated with availability and suppliers of critical ingredients. These critical ingredients are necessary for preparation of solidpropellants and explosives manufactured by the Navy. The objective of the program was to identify primary and secondary (or back-up) vendor information for these critical ingredients, and to develop suitable alternative materials if an ingredient is unavailable. In 1992 NSWC/IHDIV funded Chemical Propulsion Information Agency (CPIA) under a Technical Area Task (TAT) to expedite the task of creating a database listing critical ingredients used to manufacture Navy propellant and explosives based on known formulation quantities. Under this task CPIA provided employees that were 100 percent dedicated to the task of obtaining critical ingredient suppliers information, selecting the software and designing the interface between the computer program and the database users. TAT objectives included creating the Explosive Ingredients Source Database (EISD) for Propellant, Explosive and Pyrotechnic (PEP) critical elements. The goal was to create a readily accessible database, to provide users a quick-view summary of critical ingredient supplier's information and create a centralized archive that CPIA would update and distribute. EISD funding ended in 1996. At that time, the database entries included 53 formulations and 108 critical used to manufacture Navy propellant and explosives. CPIA turned the database tasking back over to NSWC/IHDIV to maintain and distribute at their discretion. Due to significant interest in propellant/explosives critical ingredients suppliers' status, the Propellant Development and Characterization Subcommittee (PDCS) approached the JANNAF Executive committee (EC) for authorization to continue the critical ingredient database work. In 1999, JANNAF EC approved the PDCS panel

JHU/APL conducted a series of open-air burns of small blocks (3 to 10 kg) of solid rocket motor (SRM) propellant at the Thiokol Elkton MD facility to elucidate the thermal environment under burning propellant. The propellant was TP-H-3340A for the STAR 48 motor, with a weight ratio of 71/18/11 for the ammonium perchlorate, aluminum, and HTPB binder. Combustion inhibitor applied on the blocks allowed burning on the bottom and/or sides only. Burns were conducted on sand and concrete to simulate near-launch pad surfaces, and on graphite to simulate a low-recession surface. Unique test fixturing allowed propellant self-levitation while constraining lateral motion. Optics instrumentation consisted of a longwave infrared imaging pyrometer, a midwave spectroradiometer, and a UV/visible spectroradiometer. In-situ instrumentation consisted of rod calorimeters, Gardon gauges, elevated thermocouples, flush thermocouples, a two-color pyrometer, and Knudsen cells. Witness materials consisted of yttria, ceria, alumina, tungsten, iridium, and platinum/rhodium. Objectives of the tests were to determine propellant burn characteristics such as burn rate and self-levitation, to determine heat fluxes and temperatures, and to carry out materials analyses. A summary of qualitative results: alumina coated almost all surfaces, the concrete spalled, sand moisture content matters, the propellant self-levitated, the test fixtures worked as designed, and bottom-burning propellant does not self-extinguish. A summary of quantitative results: burn rate averaged 1.15 mm/s, thermocouples peaked at 2070 C, pyrometer readings matched MWIR data at about 2400 C, the volume-averaged plume temperatures were 2300-2400 C with peaks of 2400-2600 C, and the heat fluxes peaked at 125 W/cm2. These results are higher than other researchers' measurements of top-burning propellant in chimneys, and will be used, along with Phase 3 test results, to analyze hardware response to these environments, including General

The FTIR emission spectra in the spectral range of 4,500-300 cm-1 for the solidpropellants were measured by a remote sensing FTIR system. The P-branch of fine structure of HCl fundamental band lying at 3.46 microns was used for precise combustion temperature measurement of the solidpropellant. The effect of the organic compound in the solidpropellant on the combustion temperature was discussed.

Full Text Available Binder and oxidizer decomposition play very significant role during the combustion of composite solidpropellants. Ammonium perchlorate (AP is the practical oxidizer in composite propellant formulations. Available information on binder decomposition in general and AP decomposition in particular have been collected and reviewed from the viewpoint of their application in propellants. This review may be useful in understanding the mechanism of propellant combustion.

Destruction of a solid rocket stage of a launch vehicle can create a thermal radiation hazard for an aborting crew module. This hazard was assessed for the Constellation Program (Cx) crew and launch vehicle concept. For this concept, if an abort was initiated in first stage flight, the Crew Module (CM) will separate and be pulled away from the malfunctioning launch vehicle via a Launch Abort System (LAS). Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solidpropellant, poses a threat to the CM. The harsh radiative thermal environment, caused by surrounding burning propellant debris, may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and to quantify the risk of first stage propellant debris leading to the thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Solidpropellant air-turbo-rocket (SPATR) is an air-breathing propulsion system. A numerical model of performance and characteristics analysis for SPATR was presented and the corresponding computer program was written according to the operation characteristics of SPATR. The influence on the SPATR performance at design point caused by the gas generator exit parameters and compressor pressure ratio had been computed and analyzed in detail. The off-design perform-ance of SPATR at sea level and high altitude had also been computed. The performance of thrust and specific impulse for SPATR with different solidpropellant had been compared at off-design points, and the off-design performance comparison had been made between fuel-rich and oxygen-rich. The computation results indicated that SPATR operates within wide range of Maeh number (0 ～3) and altitude (0～12 km), and SPATR possesses high specific thrust (1 200 N/(kg/s)) and high specific impulse (7000 N/ (kg/s)) when fuel-air ratio of combustor equals fuel-air ratio.

Electrically controlled extinguishable solidpropellants (ESCSP) are capable of multiple ignitions, extinguishments and throttle control by the application of electrical power. Both core and end burning no moving parts ECESP grains/motors to three inches in diameter have now been tested. Ongoing research has led to a newer family of even higher performance ECESP providing up to 10% higher performance, manufacturing ease, and significantly higher electrical conduction. The high conductivity was not found to be desirable for larger motors; however it is ideal for downward scaling to micro and pico- propulsion applications with a web thickness of less than 0.125 inch/ diameter. As a solid solution propellant, this ECESP is molecularly uniform, having no granular structure. Because of this homogeneity and workable viscosity it can be directly cast into thin layers or vacuum cast into complex geometries. Both coaxial and grain stacks have been demonstrated. Combining individual propellant coaxial grains and/or grain stacks together form three-dimensional arrays yield modular cluster thrusters. Adoption of fabless manufacturing methods and standards from the electronics industry will provide custom, highly reproducible micro-propulsion arrays and clusters at low costs. These stack and cluster thruster designs provide a small footprint saving spacecraft surface area for solar panels and/or experiments. The simplicity of these thrusters will enable their broad use on micro-pico satellites for primary propulsion, ACS and formation flying applications. Larger spacecraft may find uses for ECESP thrusters on extended booms, on-orbit refueling, pneumatic actuators, and gas generators.

This paper presents the results of detailed analyses of the images from experiments that were conducted on the formation of solid hydrogen particles in liquid helium. Solid particles of hydrogen were frozen in liquid helium, and observed with a video camera. The solid hydrogen particle sizes, their agglomerates, and the total mass of hydrogen particles were estimated. Particle sizes of 1.9 to 8 mm (0.075 to 0.315 in.) were measured. The particle agglomerate sizes and areas were measured, and the total mass of solid hydrogen was computed. A total mass of from 0.22 to 7.9 grams of hydrogen was frozen. Compaction and expansion of the agglomerate implied that the particles remain independent particles, and can be separated and controlled. These experiment image analyses are one of the first steps toward visually characterizing these particles, and allow designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Ammonium perchlorate (AP) and cyclotretamethylenetetranitramine (HMX) are two solid ingredients often used in modern solidpropellants. Although these two ingredients have very similar burning rates as monopropellants, they lead to significantly different characteristics when combined with binders to form propellants. Part of the purpose of this paper is to relate the observed combustion characteristics to the postulated flame structures and mechanisms for AP and HMX propellants that apparently lead to these similarities and differences. For AP composite, the primary diffusion flame is more energetic than the monopropellant flame, leading to an increase in burning rate over the monopropellant rate. In contrast the HMX primary diffusion flame is less energetic than the HMX monopropellant flame and ultimately leads to a propellant rate significantly less than the monopropellant rate in composite propellants. During the past decade the search for more energetic propellants and more environmentally acceptable propellants is leading to the development of propellants based on ingredients other than AP and HMX. The objective of this paper is to utilize the more familiar combustion characteristics of AP and HMX containing propellants to project the combustion characteristics of propellants made up of more advanced ingredients. The principal conclusion reached is that most advanced ingredients appear to burn by combustion mechanisms similar to HMX containing propellants rather than AP propellants.

This thesis describes the design, fabrication, and testing of a solidpropellant microthruster (SPM), which is a two-dimensional matrix of millimeter-sized rockets each capable of delivering millinewtons of thrust and millinewton-seconds of impulse to perform fine orbit and attitude corrections. The SPM is a potential payload for nanosatellites to increase spacecraft maneuverability and is constrained by strict mass, volume, and power requirements. The dimensions of the SPM in the millimeter-scale result in a number of scaling issues that need consideration such as a low Reynolds number, high heat loss, thermal and radical quenching, and incomplete combustion. The design of the SPM, engineered to address these issues, is outlined. The SPM fabrication using low-cost commercial off-the-shelf materials and standard micromachining is presented. The selection of a suitable propellant and its customization are described. Experimental results of SPM firing to demonstrate successful ignition and sustained combustion are presented for three configurations: nozzleless, sonic nozzle, and supersonic nozzle. The SPM is tested using a ballistic pendulum thrust stand. Impulse and thrust values are calculated and presented. The performance values of the SPM are found to be consistent with existing designs.

This paper presents the first, to our knowledge, direct measurement of aerosol produced by an aluminized solid rocket propellant (SRP) fire on the ground. Such fires produce aluminum oxide particles small enough to loft high into the atmosphere and disperse over a wide area. These results can be applied to spacecraft launchpad accidents that expose spacecraft to such fires; during these fires, there is concern that some of the plutonium from the spacecraft power system will be carried with the aerosols. Accident-related lofting of this material would be the net result of many contributing processes that are currently being evaluated. To resolve the complexity of fire processes, a self-consistent model of the ground-level and upper-level parts of the plume was determined by merging ground-level optical measurements of the fire with lidar measurements of the aerosol plume at height during a series of SRP fire tests that simulated propellant fire accident scenarios. On the basis of the measurements and model results, the Johns Hopkins University Applied Physics Laboratory (JHU/APL) team was able to estimate the amount of aluminum oxide (alumina) lofted into the atmosphere above the fire. The quantification of this ratio is critical for a complete understanding of accident scenarios, because contaminants are transported through the plume. This paper provides an estimate for the mass of alumina lofted into the air.

The internal ballistic effects of combined radial and circumferential grain temperature gradients are evaluated theoretically for the Space Shuttle solid rocket motors (SRMs). A simplified approach is devised for representing with closed-form mathematical expressions the temperature distribution resulting from the anticipated thermal history prior to launch. The internal ballistic effects of the gradients are established by use of a mathematical model which permits the propellant burning rate to vary circumferentially. Comparative results are presented for uniform and axisymmetric temperature distributions and the anticipated gradients based on an earlier two-dimensional analysis of the center SRM segment. The thrust imbalance potential of the booster stage is also assessed based on the difference in the thermal loading of the individual SRMs of the motor pair which may be encountered in both summer and winter environments at the launch site. Results indicate that grain temperature gradients could cause the thrust imbalance to be approximately 10% higher in the Space Shuttle than the imbalance caused by SRM manufacturing and propellant physical property variability alone.

The flame temperature of three kinds of solidpropellants was measured by passive remote sensing FTIR with the resolution of 1 cm(-1). These three kinds of solidpropellants are adulterate nano-scale metal oxide particles, adulterate normal metal oxide particles, and propellant without any adulterations. The main components of the solidpropellant are nitrocellulose and nitroglycerin. The metallic oxides, including 6 nm CuO, 56 nm Fe2O3, 16 nm NiO, and correspondingly the normal particles, were adulterated into the solidpropellants respectively. The flame temperature was calculated through the fine structure of the emission fundamental band of H2O at 2.75 microm. The results of the flame temperature of the solidpropellants adulterating nano-scale CuO, Fe2O3 and NiO are 3089, 3193 and 3183 K, respectively. The temperatures of the three kinds of solidpropellants were compared, and it was shown that there is no obvious difference in the flame temperature among the three kinds of solidpropellants.

An analysis of launch vehicle Gross Liftoff Weight (GLOW) using high energy density atomic propellants with solid particle feed systems was conducted. The analyses covered several propellant combinations, including atoms of aluminum (Al), boron (B). carbon (C), and hydrogen (H) stored in a solid cryogenic particle, with a cryogenic liquid as the carrier fluid. Several different weight percents (wt%) for the liquid carrier were investigated and the gross lift off weight (GLOW) of the vehicles using the solid particle feed systems were compared with a conventional 02/H2 propellant vehicle. The potential benefits and effects of feed systems using solid particles in a liquid cryogenic fluid are discussed.

A new cold flow test facility was designed and constructed at NASA Marshall Space Flight Center for the purpose of characterizing the flow field in the port and nozzle of solidpropellant rocket motors (SRM's). A National Advisory Committee was established to include representatives from industry, government agencies, and universities to guide the establishment of design and instrumentation requirements for the new facility. This facility design includes the basic components of air storage tanks, heater, submicron filter, quiet control valve, venturi, model inlet plenum chamber, solid rocket motor (SRM) model, exhaust diffuser, and exhaust silencer. The facility was designed to accommodate a wide range of motor types and sizes from small tactical motors to large space launch boosters. This facility has the unique capability of testing ten percent scale models of large boosters such as the new Advanced Solid Rocket Motor (ASRM), at full scale motor Reynolds numbers. Previous investigators have established the validity of studying basic features of solid rocket motor development programs include the acquisition of data to (1) directly evaluate and optimize the design configuration of the propellant grain, insulation, and nozzle; and (2) provide data for validation of the computational fluid dynamics, (CFD), analysis codes and the performance analysis codes. A facility checkout model was designed, constructed, and utilized to evaluate the performance characteristics of the new facility. This model consists of a cylindrical chamber and converging/diverging nozzle with appropriate manifolding to connect it to the facility air supply. It was designed using chamber and nozzle dimensions to simulate the flow in a 10 percent scale model of the ASRM. The checkout model was recently tested over the entire range of facility flow conditions which include flow rates from 9.07 to 145 kg/sec (20 to 320 Ibm/sec) and supply pressure from 5.17 x 10 exp 5 to 8.27 x 10 exp 6 Pa. The

The CAEPE, the French Centre of Propellers and Engines Completion and Testing is in charge of the static tests of solidpropellant fuelled propellers. In order to determine the schedule of firing permissions, predictive means are used to predict the environmental impact of propellers firing. Calculation and simulation codes are used to build maps of acoustic nuisance and acid fallout. These codes, which use in-situ meteorological radio sounding data, were progressively adjusted during testing of engines with different sizes (up to the Ariane 5 P230 auxiliary propeller). In this presentation, the authors focus on a model derived from the G.A. Briggs` model for the ascension of high temperature effluents. The correct simulation of acid rains requires a good description of this phenomenon. Concerning the other aspects of the codes, the main references are given. (J.S.) 15 refs.

The study on spinning solid rocket motor (SRM) which used as power plant of twice throwing structure of aerial submunition was introduced. This kind of SRM which with the structure of tangential multi-nozzle consists of a combustion chamber, propellant charge, 4 tangential nozzles, ignition device, etc. Grain design, structure design and prediction of interior ballistic performance were described, and problem which need mainly considered in design were analyzed comprehensively. Finally, in order to research working performance of the SRM, measure pressure-time curve and its speed, static test and dynamic test were conducted respectively. And then calculated values and experimental data were compared and analyzed. The results indicate that the designed motor operates normally, and the stable performance of interior ballistic meet demands. And experimental results have the guidance meaning for the pre-research design of SRM.

Full Text Available A family of propellants based on a low cost hydroxy terminated binder has been developed and proved in large size motors. It can meet the requirements of Apogee motors as well as large boosters. The system offers advantages comparable with HTPB propellants in terms of high ballistic performance, stringent mechanical properties, ease and reliability of cure even at ambient conditions and high storage stability. The near-Newtonian flow behaviour, simplicity and processing characteristics of this saturated binder propellant are particularly note-worthy.

Full Text Available In this study, three dimensional modelling of extrusion forming of a double base solid rocket propellant is performed on Ansys® finite element simulation package. For the purpose of initial model construction and later comparisons with elastoviscoplastik model, the solidpropellant is assumed to obey the elastic-plastic material response during the direct extrusion process. Taking into account the contact surface behavior with Coulomb friction and geometric and material nonlinearities, an incremental large large strain solution methodology has been adapted in the simulation. The hydrostatic pressure, stress, strain, and displacement values during extrusion of the solid rocket propellant are obtained from the simulation.

Measurements of gaseous hydrogen chloride (HCl) and particulate aluminum oxide (Al2O3) were made during penetrations of five Space Shuttle exhaust clouds and one static ground test firing of a shuttle booster. Instrumented aircraft were used to penetrate exhaust clouds and to measure and/or collect samples of exhaust for subsequent analyses. The focus was on the primary solid rocket motor exhaust products, HCl and Al2O3, from the Space Shuttle's solid boosters. Time-dependent behavior of HCl was determined for the exhaust clouds. Composition, morphology, surface chemistry, and particle size distributions were determined for the exhausted Al2O3. Results determined for the exhaust cloud from the static test firing were complicated by having large amounts of entrained alkaline ground debris (soil) in the lofted cloud. The entrained debris may have contributed to neutralization of in-cloud HCl.

Using the theory of gas dynamics and heat transfer from a turbulent gas flow to the burning surface of propellant along a permeable wall, an explicit expression is derived to predict the burning rate of the solidpropellant with crossflow. Results of the calculation have been compared with experimental data and proved to be correct.

Full Text Available Combustion instability in solidpropellant rocket motor is numerically simulated by implementing propellant response function with quasi steady homogeneous one dimensional formulation. The convolution integral of propellant response with pressure history is implemented through a user defined function in commercial computational fluid dynamics software. The methodology is validated against literature reported motor test and other simulation results. Computed amplitude of pressure fluctuations compare closely with the literarture data. The growth rate of pressure oscillations of a cylindrical grain solid rocket motor is determined for different response functions at the fundamental longitudinal frequency. It is observed that for response function more than a critical value, the motor exhibits exponential growth rate of pressure oscillations.

The NASA KSC VAB was built to process Apollo launchers in the 1960's, and later adapted to process Space Shuttles. The VAB has served as a place to assemble solid rocket motors (5RM) and mate them to the vehicle's external fuel tank and Orbiter before rollout to the launch pad. As Space Shuttle is phased out, and new launchers are developed, the VAB may again be adapted to process these new launchers. Current launch vehicle designs call for continued and perhaps increased use of SRM segments; hence, the safe separation distances are in the process of being re-calculated. Cognizant NASA personnel and the solid rocket contractor have revisited the above VAB QD considerations and suggest that it may be revised to allow a greater number of motor segments within the VAB. This revision assumes that an inadvertent ignition of one SRM stack in its High Bay need not cause immediate and complete involvement of boosters that are part of a vehicle in adjacent High Bay. To support this assumption, NASA and contractor personnel proposed a strawman test approach for obtaining subscale data that may be used to develop phenomenological insight and to develop confidence in an analysis model for later use on full-scale situations. A team of subject matter experts in safety and siting of propellants and explosives were assembled to review the subscale test approach and provide options to NASA. Upon deliberations regarding the various options, the team arrived at some preliminary recommendations for NASA.

The current MK 66 2.75 inch Rocket Motor double-base propellant contains the lead-based ballistic modifier LC-12-15 to achieve the desired plateau and mesa burning rate characteristics. The use of lead compounds poses a concern for the environment and for personal safety due to the metal's toxic nature when introduced into the atmosphere by propellant manufacture, rocket motor firing, and disposal. Copper beta-resorcylate (copper 2,4-di-hydroxy-benzoate) was successfully used in propellant as a simple modifier in the mid 1970's. This and other compounds have also been mixed with lead salts to obtain more beneficial ballistic results. Synthesized complexes of lead and copper compounds soon replaced the mixtures. The complexes incorporate the lead, copper lack of organic liquids, which allows for easier propellant processing. About ten years ago, the Indian Head Division, Naval Surface Warfare Center (NSWC), initiated an effort to develop a lead-free propellant for use in missile systems. Several lead-free propellant candidate formulations were developed. About five years ago, NSWC, in conjunction with Alliant Techsystems, Radford Army Ammunition Plant, continued ballistic modifier investigations. A four component ballistic modifier system without lead for double-base propellants that provide adequate plateau and mesa burn rate characteristics was developed and patented. The ballistic modifier's system contains bismuth subsalicylate, 1.5 percent; copper salicylate, 1.0 percent, copper stannate, 0.77 percent; and carbon black, 0.1 percent. Action time and impulse data obtained through multiple static firings indicate that the new lead-free double-base propellant, while not a match for NOSIH-AA-2, will be a very suitable replacement in the 2.75 inch Rocket Motor. Accelerated aging of the double-base propellant containing the lead-free ballistic modifier showed that it had a much higher rate of stabilizer depletion than the AA-2. A comprehensive study showed that an

In order to improve the safety of high-energy solidpropellants, a study is carried out for the effects of damage on the combustion of the NEPE (Nitrate Ester Plasticized Polyether) propellant. The study includes: (1) to introduce damage into the propellants by means of a large-scale drop-weight apparatus; (2) to observe microstructural variations of the propellant with a scanning electron microscope (SEM) and then to characterize the damage with density measurements; (3) to investigate thermal decomposition; (4) to carry out closed-bomb tests. The NEPE propellant can be considered as a viscoelastic material. The matrices of damaged samples are severely degraded, but the particles are not. The results of the thermal decomposition and closed-bomb tests show that the microstructural damage in the propellant affects its decomposition and burn rate.

Solidpropellants, sand-asphalt concrete and hard plastics showed rate sensitive mechanical behavior which, in addition, indicated that these materials have a permanent memory of the strain (or loading) path by which their present state was attained. A constitutive equation was formulated in general three dimensional tensorial form by means of irreversible thermodynamics. By using a very simple analytical form, it was shown that the mechanical behavior of solidpropellants and sand-asphalt concrete can be readily described.

Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

Multiple flame models derived for simple composite propellants are extended to describe the combustion of propellants containing multimodal particle sizes, mixed oxidizers and monopropellant binders. Models combining the component contributions to propellant surface structure, flame structure and energy distribution are based in part upon experimental observations and in part upon hypotheses constrained to provide reasonable agreement with measured burning rate characteristics. The methods employed consist of superposition, interaction and iteration. The computerized model is applied to explain the effects of multiple ingredients and to discuss burning rate tailoring problems of current interest.

The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site.

Full Text Available The dynamic mechanical properties of different binders and corresponding propellants are studied in terms of storage modulus and loss tangent. The binders investigated are HTPB, CTPB, PBAN, HEF-20 and ISRO polyol. The viscoelastic behaviour is investigated using Rheovibron viscoelastometer at 35 Hz covering a wide temperature range (-100 degree centigrade to 100 degree centigrade. The properties of the binder and corresponding propellant are compared in terms of parameters, tan delta/sub max/, T/sub g/ and the trend of their master relaxation modulus curves. It is found that polybutadiene binders exhibit lowest T/sub g/ (around -60 degree centigrade and ISRO polyol the highest (near -20 degree centigrade. The propellants have higher moduli than the binders at any temperature. The master relaxation modulus curve is influenced by the type of propellant.

Aerojet in the mid 1960s, under contract to NASA, built and static hot fire tested the largest solid rocket motor (SRM) in history for the purpose of demonstrating the feasibility of utilizing large SRMs for space exploration. This program successfully fabricated two high strength steel chambers, loaded each with approximately 1,68 million pounds of propellant, and static test fired these giants with their nozzles up from an underground silo located adjacent to the Florida everglades. Maximum thrust and total impulse in excess of 5,000,000 lbf and 3,470,000,000 lbf-sec were achieved. Flames from the second firing, conducted at night, were seen over eighty miles away. For comparative purposes: the thrust developed was nearly 100 times that of a Minuteman III second stage and the 260 in.-dia cross-section was over 3 times that of the Space Shuttle SRM.

Full Text Available This paper reports the evaluation of four different energetic plasticisers, viz., glycidyl azidepolymer (GAP, MW = 390, 1,5-diazido-3-nitrazapentane (DANPE, ethylene-glycol-bis-azido-acetate (EGBAA and N-n-butyl-N-(2 nitroxyethyl nitramine (n-Bu-NENA separately into highenergy gun propellant containing 28 per cent NC (13.1 N %, 65 per cent RDX, 6 per cent di-octyl-phthahate (DOP and 1 per cent carbamite. Four different propellant compositions based on theenergetic plasticiser have been formulated separately with the replacement of non-energeticplasticiser, DOP. The propellants were processed by standard solvent method and evaluatedexperimentally along with the control composition to determine the ballistic parameters, cal-val,sensitivity, thermal characterisation, thermal stability and mechanical properties. The performanceof the propellants containing the energetic plasticiser has been compared with that of thecontrol composition containing the non-energetic plasticiser, DOP so as to assess the suitabilityof the energetic plasticiser for the futuristic gun propellant formulations. It has been found outthat n-Bu-NENA is the superior plasticiser among the four energetic plasticisers evaluated inthis study.Defence Science Journal, 2008, 58(1, pp.86-93, DOI:http://dx.doi.org/10.14429/dsj.58.1627

The Advanced Solid Rocket Motor is a new design for the Space Shuttle Solid Rocket Booster. The new design will provide more thrust and more payload capability, as well as incorporating many design improvements in all facets of the design and manufacturing process. A 48-inch (diameter) test motor program is part of the ASRM development program. This program has multiple purposes for testing of propellent, insulation, nozzle characteristics, etc. An overview of the evolution of the 48-inch ASRM test motor ignition system which culminated with the implementation of a laser ignition system is presented. The laser system requirements, development, and operation configuration are reviewed in detail.

Differential pulse polarography (DPP) and square wave voltammetry (SWV) were investigated, in order to know the stability of solidpropellants which contain diphenylamine. The simultaneous determination of N-nitrousdiphenylamine (NnDPA), 2-nitrodiphenylamine (2nDPA), 4-nitrodiphenylamine (4nDPA) and 2,4-dinitrodiphenylamine (2,4dnDPA) by DPP and SWV was proposed due to these nitro-derivatives appear during the stabilisation process from degradation of diphenylamine (DPA) used as stabiliser in propellant compositions. The proposed methods were successfully applied to the simple base solidpropellant (with nitrocellulose as the only active component), with its stabiliser DPA. In all cases and with both the techniques, detection limits propellants accurately, faster and cheaper than LC methods.

The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site. Sites being considered for the new facilities include John C. Stennis Space Center, Hancock County, Mississippi; the Yellow Creek site in Tishomingo County, Mississippi, which is currently in the custody and control of the Tennessee Valley Authority; and John F. Kennedy Space Center, Brevard County, Florida. TVA proposes to transfer its site to the custody and control of NASA if it is the selected site. All facilities need not be located at the same site. Existing facilities which may provide support for the program include Michoud Assembly Facility, New Orleans Parish, Louisiana; and Slidell Computer Center, St. Tammany Parish, Louisiana. NASA's preferred production location is the Yellow Creek site, and the preferred test location is the Stennis Space Center.

Full Text Available Copper chromites are well known as burning rate catalysts for the combustion of composite solidpropellants, used as a source of energy for rocket propulsion. The propellant burning rate depends upon the catalyst characteristics such as chemical composition and specific surface area. In this work, copper chromite samples from different suppliers were characterized by chemical analysis, FT-IR spectroscopy and by surface area measurement (BET. The samples were then evaluated as burning rate catalyst in a typical composite propellant formulation based on HTPB binder, ammonium perchlorate and aluminum. The obtained surface area values are very close to those informed by the catalyst suppliers. The propellant processing as well as its mechanical properties were not substantially affected by the type of catalyst. Some copper chromite catalysts caused an increase in the propellant burning rate in comparison to the iron oxide catalyst. The results show that in addition to the surface area, other parameters like chemical composition, crystalline structure and the presence of impurities might be affecting the catalyst performance. All evaluated copper chromite samples may be used as burning rate catalyst in composite solidpropellant formulations, with slight advantages for the SX14, Cu-0202P and Cu-1800P samples, which led to the highest burning rate propellants.

Pre-World War II Soviet solid-propellant rocket technology is reviewed. Research and development regarding solid composite preparations of pyroxyline TNT powder is described, as well as early work on rocket loading calculations, problems of flight stability, and aircraft rocket launching and ground rocket launching capabilities.

A standard quench bomb (QB) - widely used to characterize condensed phase from metalized solidpropellant combustion - is studied in detail. Experimental and numerical investigations proved that collected particles are mostly unburned aluminum (Al) agglomerates despite large quenching distances. Particles are actually found to quench early as propellant surface is swept by inert pressurant. Further improvements of the QB are proposed which allow measuring both Al agglomerates and alumina residue with the same setup. Finally, the results obtained on a typical aluminized ammonium perchlorate (AP) / hydroxyl-terminated polybutadiene (HTPB) propellant are briefly discussed.

Full Text Available Studies on the pyrolysis leading to ignition of polystyrene (PS/ammonium perchlorate (AP, polyvinyl chloride (PVC/AP and polyphenol formaldehyde (PPF/AP propellants revealed that the activation energy for the ignition strongly depends upon the binder. Double base propellants exhibit an empirical relationship between the ignition delay (~ and the oxidiser concentration; when the same equation is applied to composite solidpropellants, the plot of In ~ vs In C (C=Concentration, % yields a straight line with a knee corresponding to 65-70 per cent AP above which the dependence on ~ becomes less sensitive.

An analytical technique for implementing simultaneously the temperature, dynamic strain, real modulus, and frequency properties of solidpropellant in an unsymmetrical vibrating ring mode is presented. All dynamic parameters and sources are defined for a free vibrating ring-grain structure with initial displacement and related to a forced vibrating system to determine the change in real modulus. Propellant test data application is discussed. The technique was developed to determine the aft attach ring stiffness of the shuttle booster at lift-off.

Storage of solidpropellants in either a dry or a vacuum environment causes a significantly greater increase in the propellants' modulus and maximum tensile strength than does ambient storage. It is postulated that these physical property changes can be attributed to the effect trace amount of moisture has on the bond between the propellants' binder and oxidizer.

Full Text Available There is little doubt that explosives had their origin in warfare. In the armed conflict between groups of individuals or of states, where each sought and still seeks to impose its will upon the other by force, it was inevitable that arms should grow and flourish. The sling, the bow and arrow, the sword and firearm typify evolution in warfare weapons. Bs a means of propelling missiles, the gun and gun powder were thought of. The history of explosives, therefore, may be said to begin with black powder.

This project was started in the interest of supplementing existing data on additives to composite solidpropellants. The study on the addition of iron and aluminum nanoparticles to composite AP/HTPB propellants was conducted at the Combustion and Energy Systems Laboratory at RPI in the new strand-burner experiment setup. For this study, a large literature review was conducted on history of solidpropellant combustion modeling and the empirical results of tests on binders, plasticizers, AP particle size, and additives. The study focused on the addition of nano-scale aluminum and iron in small concentrations to AP/HTPB solidpropellants with an average AP particle size of 200 microns. Replacing 1% of the propellant's AP with 40-60 nm aluminum particles produced no change in combustive behavior. The addition of 1% 60-80 nm iron particles produced a significant increase in burn rate, although the increase was lesser at higher pressures. These results are summarized in Table 2. The increase in the burn rate at all pressures due to the addition of iron nanoparticles warranted further study on the effect of concentration of iron. Tests conducted at 10 atm showed that the mean regression rate varied with iron concentration, peaking at 1% and 3%. Regardless of the iron concentration, the regression rate was higher than the baseline AP/HTPB propellants. These results are summarized in Table 3.

Full Text Available The sensitivity of flaw detection with x-ray radiographic methods is investigated here qualitatively in case of cast double base and cast composite propellants and for air pockets it is found to be 1.5 and 0.9 percent of the web respectively. General guidelines for the inspection of sustainer charges have also been laid down.

Specially designed sensor is attached to or imbedded in propellant. When sensor is driven into vibration, it moves with a phase lag directly proportional to internal friction or loss coefficent. Resonance frequency of the system is related to Young's modulus. Modulus or internal friction can be monitored over long period of time.

This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems.This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for compressible flows only, was modified to work, proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of references available in this area. Here an existing code for compressible flow was used and primitive variables,the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to permit the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and extemal flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theoretical results and with other codes already validated and well accepted by the CFD community.

and then create a 30 model in SolidWorks . The MATLAB G Ul ultimately results in a mostly automated process that is simple to use for individuals who...are unfamiliar with command prompt programs and SolidWorks modeling. Also incorporated into BEARCONTROL is the program NREL AirFoil Noise (NAFNOISE...propeller design, QPROP for the motor and propeller performance analysis, NAFNoise for the propeller noise prediction, and SolidWorks for the solid

The starting characteristics of thermodynamic undersea vehicle systems are determined by the geometry, size and combustion area of solidpropellants, which directly effect liquid propellant pipeline design. It is necessary to establish accurate burning models for solidpropellants. Based on combustion models using powder tings and two different solid ignition grains, namely star-shaped ignition grains and stuffed ignition grains, a mathematic model of the ignition process of the propulsion system was built.With the help of Matlah, a series of calculations were made to determine the effects of different grains on ignition characteristics. The results show that stuffed ignition grain is best suited to be the ignition grain of a thermodynamic undersea vehicle system.

Full Text Available A particle swarm/pattern search hybrid optimizer was used to drive a solid rocket motor modeling code to an optimal solution. The solidmotor code models tapered motor geometries using analytical burn back methods by slicing the grain into thin sections along the axial direction. Grains with circular perforated stars, wagon wheels, and dog bones can be considered and multiple tapered sections can be constructed. The hybrid approach to optimization is capable of exploring large areas of the solution space through particle swarming, but is also able to climb “hills” of optimality through gradient based pattern searching. A preliminary method for designing tapered internal geometry as well as tapered outer mold-line geometry is presented. A total of four optimization cases were performed. The first two case studies examines designing motors to match a given regressive-progressive-regressive burn profile. The third case study studies designing a neutrally burning right circular perforated grain (utilizing inner and external geometry tapering. The final case study studies designing a linearly regressive burning profile for right circular perforated (tapered grains.

The infrared radiation signature of the plume from solidpropellants with different energy characteristics is not the same.Three kinds of double-base propellants of different energy characteristics are chosen to measure the infrared spectral radiance from 1000cm-1 to 4500 cm-1 of their plumes.The radiative spectrum is obtained in the tests.The experimental results indicate that the infrared radiation of the plume is determined by the energy characteristics of the propellant.The radiative transfer calculation models of the exhaust plume for the solidpropellants are established.By including the chemical reaction source term and the radiation source term into the energy equation,the plume field and the radiative transfer are solved in a coupled way.The calculated results are consistent with the experimental data,so the reliability of the models is confirmed.The temperature distribution and the extent of the afterburning of the plume are distinct for the propellants of different energy characteristics,therefore the plume radiation varies for different propellants.The temperature of the fluid cell in the plume will increase or decrease to some extent by the influence of the radiation term.

Full Text Available This paper reviews the research work carried out in the field of metallised fuel-rich propellants (FRP. Limitations and merits of various potential metals (Al, Mg, B, Be, or Zr as a component of FRP are discussed. The paper also includes a discussion on the combustion mechanism of metallised propellants, including problem areas and probable remedial measures. Zirconium and Ti appear to have potential to offer FRP with efficient combustion. Ideal performance is not achieved with current systems based on Al and B and further work is needed to develop FRP having all three desirable attributes, viz., ease of ignition, stable combustion and high specific impulse (I/sub sp/ in a single composition.

We examine random packs of discs or spheres, models for ammonium-perchlorate-in-binder propellants, and discuss their average properties. An analytical strategy is described for calculating the mean or effective heat conduction coefficient in terms of the heat conduction coefficients of the individual components, and the results are verified by comparison with those of direct numerical simulations (dns) for both 2-D (disc) and 3-D (sphere) packs across which a temperature difference is applied. Similarly, when the surface regression speed of each component is related to the surface temperature via a simple Arrhenius law, an analytical strategy is developed for calculating an effective Arrhenius law for the combination, and these results are verified using dns in which a uniform heat flux is applied to the pack surface, causing it to regress. These results are needed for homogenization strategies necessary for fully integrated 2-D or 3-D simulations of heterogeneous propellant combustion.

Deflagration to Detonation (DDT) was predicted to occur in porous beds of high-energy solidpropellants by solving the unsteady fluid mechanical convective heat transfer from hot gas products, obtained from the rapid burning at high pressures, provides the impetus to develop a narrow combustion zone and a resulting strong shock. A parametric study clearly indicates that DDT occurs only when a combination of the solids loading fraction, the burning rate constants, the propellant chemical energy, and the particle size provide for critical energy and gas release to support a detonation wave. Predictions for the run-up length to detonation as a function of these parameters are presented.

A progress report is given on a research project to use the microwave Doppler velocimeter technique to measure the combustion response to an oscillating thermal radiation source (CO2 laser). The test technique and supporting analyses are described, and the results are presented for an initial test series on the nonmetallized, composite propellant, Naval Weapons Center formulation A-13. It is concluded that in-depth transmission of radiant heat flux is not a factor at the CO2 laser wave length.

Full Text Available A 30 mm electrothermal-chemical (ETC gun experimental system is employed to research the burning rate characteristics of 4/7 high-nitrogen solidpropellant. Enhanced gas generation rates (EGGR of propellants during and after electrical discharges are verified in the experiments. A modified 0D internal ballistic model is established to simulate the ETC launch. According to the measured pressure and electrical parameters, a transient burning rate law including the influence of EGGR coefficient by electric power and pressure gradient (dp/dt is added into the model. The EGGR coefficient of 4/7 high-nitrogen solidpropellant is equal to 0.005 MW−1. Both simulated breech pressure and projectile muzzle velocity accord with the experimental results well. Compared with Woodley's modified burning rate law, the breech pressure curves acquired by the transient burning rate law are more consistent with test results. Based on the parameters calculated in the model, the relationship among propellant burning rate, pressure gradient (dp/dt and electric power is analyzed. Depending on the transient burning rate law and experimental data, the burning of solidpropellant under the condition of plasma is described more accurately.

A 30 mm electrothermal-chemical (ETC) gun experimental system is employed to research the burning rate characteristics of 4/7 high-nitrogen solidpropellant. Enhanced gas generation rates (EGGR) of propellants during and after electrical discharges are verified in the experiments. A modified 0D internal ballistic model is established to simulate the ETC launch. According to the measured pressure and electrical parameters, a transient burning rate law including the influence of EGGR coefficient by electric power and pressure gradient (dp/dt) is added into the model. The EGGR coefficient of 4/7 high-nitrogen solidpropellant is equal to 0.005 MW−1. Both simulated breech pressure and projectile muzzle velocity accord with the experimental results well. Compared with Woodley’s modified burning rate law, the breech pressure curves acquired by the transient burning rate law are more consistent with test results. Based on the parameters calculated in the model, the relationship among propellant burning rate, pressure gradient (dp/dt) and electric power is analyzed. Depending on the transient burning rate law and experimental data, the burning of solidpropellant under the condition of plasma is described more accurately.

Full Text Available On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

Full Text Available The shelf life of a composite solidpropellant is one of the critical aspects for the usage of solidpropellants. To assess the ageing behavior of the composite solidpropellant, the activation energy is a key parameter. The activation energy is determined by analysis of visco-elastic response of the composite solidpropellant when subjected to sinusoidal excitation. In the present study, dynamic mechanical analyzer was used to characterize six different types of propellants based on hydroxyl terminated polybutadiene, aluminium, ammonium perchlorate cured with toluene diisocyanate having burning rates varying from 5 mm/s to 25 mm/s at 7000 kPa. Each propellant sample was given a multi-frequency strain of 0.01 percent at three discrete frequencies (3.5 Hz, 11 Hz, 35 Hz in the temperature range -80 °C to + 80 °C. It was observed that all the propellants have shown two relaxation events (α- and β- transition in the temperature range -80 °C to +80 °C. The α-transition was observed between -66 °C and -51 °C and β-transition between 7 °C and 44 °C for the propellants studied. The activation energy for both transitions was determined by Arrhenius plot from dynamic properties measured at different frequencies and also by time temperature superposition principle using Williams-Landel-Ferry and Arrhenius temperature dependence equations. The data reveal that the activation energy corresponding to α-transition varies from 90 kJ/mol to 125 kJ/mol for R-value between 0.7 to 0.9 while for β-transition the values are from 75 kJ/mol to 92 kJ/mol. The activation energy corresponding to β-transition may be used to predict the useful life of solidpropellant.Defence Science Journal, 2014, 64(2, pp. 173-178. DOI: http://dx.doi.org/10.14429/dsj.64.3818

The temperature in an aluminized propellant is determined as a function of height and plume depth from diatomic AlO and thermal emission spectra. Higher in the plume, 305 and 508 mm from the burning surface, measured AlO emission spectra show an average temperature with 1σ errors of 2980 ± 80 K. Lower in the plume, 152 mm from the burning surface, an average AlO emission temperature of 2450 ± 100 K is inferred. The thermal emission analysis yields higher temperatures when using constant emissivity. Particle size effects along the plume are investigated using wavelength-dependent emissivity models.

The spectral and temporal details of the flames of a series of ammonium perchlorate-polyurethane propellants during both unstable and stable combustion were observed experimentally. A 400-scan per second optical spectrometer operating in the middle infrared region was used. During unstable combustion at low ratios of chamber free volume to nozzle throat area, three different frequencies were observed simultaneously. These were attributable to at least two mechanisms. During stable combustion periodic fluctuations in flame temperature and composition were also observed. Some aspects of theory of bulk mode instability were confirmed, but the assumptions of constant flame temperature and constant composition were found to be inaccurate.

Full Text Available Nano-catalysts containing copper–cobalt oxides (Cu–Co–O have been synthesized by the citric acid (CA complexing method. Copper (II nitrate and Cobalt (II nitrate were employed in different molar ratios as the starting reactants to prepare three types of nano-catalysts. Well crystalline nano-catalysts were produced after a period of 3 hours by the calcination of CA–Cu–Co–O precursors at 550 °C. The phase morphologies and crystal composition of synthesized nano-catalysts were examined using Scanning Electron Microscope (SEM, Energy Dispersive Spectroscopy (EDS and Fourier Transform Infrared Spectroscopy (FTIR methods. The particle size of nano-catalysts was observed in the range of 90 nm–200 nm. The prepared nano-catalysts were used to formulate propellant samples of various compositions which showed high reactivity toward the combustion of HTPB/AP-based composite solidpropellants. The catalytic effects on the decomposition of propellant samples were found to be significant at higher temperatures. The combustion characteristics of composite solidpropellants were significantly improved by the incorporation of nano-catalysts. Out of the three catalysts studied in the present work, CuCo-I was found to be the better catalyst in regard to thermal decomposition and burning nature of composite solidpropellants. The improved performance of composite solidpropellant can be attributed to the high crystallinity, low agglomeration and lowering the decomposition temperature of oxidizer by the addition of CuCo-I nano-catalyst.

303913 11. TITLE (Include Security Clasification ) Mutagenic Potential of DIGL-RP SolidPropellant in the Ames Salmonella/Mammalian Microsome Mutagenicity...mg protein /plate. After all the ingredients were added, the top agar was mixed, then overlaid on MGA plates. These plates contained 2% glucose and

Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solidpropellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

Since combustion is an easy way to achieve large quantities of energy from a small volume, we developed a MEMS based solidpropellant microthruster array for small spacecraft and micro-air-vehicle applications. A thruster is composed of a fuel chamber layer, a top-side igniter with a micromachined nozzle in the same silicon layer. Layers are assembled by adhesive bonding to give final MEMS array. The thrust force is generated by the combustion of propellant stored in a few millimeter cube chamber. The micro-igniter is a polysilicon resistor deposited on a low stress SiO2/SiNx thin membrane to ensure a good heat transfer to the propellant and thus a low electric power consumption. A large range of thrust force is obtained simply by varying chamber and nozzle geometry parameters in one step of Deep Reactive Ion Etching (DRIE). Experimental tests of ignition and combustion employing home made (DB+x%BP) propellant composed of a DoubleBase and Black-Powder. A temperature of 250 °C, enough to propellant initiation, is reached for 40 mW of electric power. A combustion rate of about 3.4 mm/s is measured for DB+20%BP propellant and thrust ranges between 0.1 and 3,5 mN are obtained for BP ratio between 10% and 30% using a microthruster of 100 μm of throat wide.

Experimental results concerning the transient burning-rate augmentation of a 16% aluminum polybutadiene acrylic acid (PBAA) propellant burned in a 2-in. web motor at pressure levels from 300 to 1200 psia with centrifugal accelerations from 0 to 140 g. The orientation of the acceleration vector was varied to determine its effect on the transient burning rate. The burning-rate augmentation was strongly dependent on (1) acceleration level, (2) propellant distance burned (or burn time), and (3) orientation of the acceleration vector with respect to the burning surface. This transient rate augmentation resulted from the retention of molten metallic residue on the burning surface by the normal acceleration loading. The presence of the residue altered the combustion zone heat transfer and caused increased localized burning rates, as evidenced by the pitted propellant surfaces that were observed from extinction tests conducted at various acceleration levels.

When an acoustic wave is present in a solidpropellant combustion environment, the mass flux from the combustion zone oscillates at the same frequency as the acoustics. The acoustic wave is either amplified or attenuated by the response of the combustion to the acoustic disturbance. When the acoustic wave is amplified, this process is called combustion instability. The amplification is quantitatively measured by a response function. The ability to predict combustion stability for a solidpropellant formulation is essential to the formulator to prevent or minimize the effects of instabilities, such as an oscillatory thrust. Unfortunately, the prediction of response values for a particular propellant remains a technical challenge. Most predictions of the response of propellants are based on test data, but there are a number of questions about the reliability of the standard test method, the T-burner. Alternate methods have been developed to measure the response of a propellant, including the ultrasound burner, the magnetic flowmeter and the rotating valve burner, but there are still inconsistencies between the results obtained by these different methods. Aside from the experimental differences, the values of the pressure-coupled responses obtained by different researchers are often compared erroneously, for the simple reason that inconsistencies in the definitions of the responses and admittances are not considered. The use of different definitions has led to substantial confusion since the first theoretical treatments of the problem by Hart and McClure in 1959. The definitions and relations derived here seek to alleviate this problem.

Full Text Available Mechanical property evaluation of composite solid rocket propellants is used as a quick quality control tool for propellant development and production. However, stress-strain curves from uni-axial tensile testing can be utilised to assess the shelf-life of propellants also. Composite propellants (CP of two varieties cartridge-loaded (CLCP and case-bonded (CBCP are utilized in rocket and missile applications. Both classes of propellants were evaluated for mechanical properties namely tensile strength, modulus and percentage elongation using specimens conforming to ASTM D638 type IV at different ageing time. Both classes of propellants show almost identical variation in various mechanical properties with time. Tensile strength increases with time for both classes of propellants and percentage elongation reduces. Initial modulus is also found to decrease with time. Tensile strength is taken as degradation criteria and it is observed that CLCP has slower degradation rate than CBCP. This is because of two facts–(i higher initial tensile strength of CLCP (1.39 MPa compared to CBCP (0.665 MPa and (ii lower degradation rate of CLCP (0.0014 MPa/day with respect to CBCP (0.0025 MPa/day. For the studied composite propellants, a degradation criterion in the form of percentage change in tensile strength is evaluated and shelf life for different degradation criteria is tabulated for quick reference.Defence Science Journal, 2012, 62(2, pp.90-94, DOI:http://dx.doi.org/10.14429/dsj.62.773

Full Text Available In this paper a comparative analysis on the energetics of ester type plasticizers such as dioctyl adipate (DOA, dioctyl phthalate (DOP,dibutyl sebacate (DBS, isodecyl pelargonate (IDP, trioctylphosphate (TOF, diethyl phthalate (DEP, tricresyl phosphate (TCPand dibutyl phthalate (DBP and hydrocarbon type plasticizers such as polybutene (PB, spindle oil, naphthenic oil, polymer extender oil(PEO and poly isobutylene (PIB and the impact of some of the plasticizers on the work ability, pot life and mechanical properties of propellants based on two selected polymeric binders namely polybutadiene-acrylic acid-acrylonitrile (PBAN ter polymer andhydroxyl terminated polybutadiene (HTPB have been reported. The compatibility of all the plasticizers on HTPB binder was also studied at different concentration levels and temperatures using Brookfield viscometer and reported. The mechanism of plasticization is also reviewed.

results of two solutions that tesi the ability of the LW+H+ACM to trect such problems are discussed. In both cases, the initial Pressure distur$.;nce...piezoelectric transducers. Thermal shielding of the pressure transducers was accomplished using vulcanized rubber (RTV). Based on shock i,,be calibration tests...potted with vulcanized rubber (RTV) which is utilized to provide thermal shielding 24 45. 16.0 AFT CLOSURE WI0.45 250. for the pressure transducers

variables, such as winding speed and tension, will affect the resin properties and, hence, the laminate properties. . 3-dimensionsl F1nite Elmaent In...accurately characterize the cbndi- tions in the composite, including the fiber-resin interaction and the resin propertie* in the laminate Eli]. The resin ... properties in the laminate may differ significantly from test coupons because of processing variations and conditions that cannot be duplicated in a

Full Text Available Finite element analysis of case-bonded solidpropellants in finocyl port configuration hasbeen carried out using finite element method. The parametric studies have also been conductedfor loading conditions, material properties, and geometrical configurations. The results arepresented in the form of a universal power law, which can be utilised for primary assessmentof peak strain in any finocyl port propellant configuration without using finite element software.This eliminates dependence on finite element software for structural integrity analysis of solidpropellants in finocyl port configuration under port pressurisation. The results obtained by finiteelement analysis and power law are in close agreement.

Propellant characteristics for solid rocket motors were not completely determined for its use as a processing variable in today's production facilities. A major effort to determine propellant characteristics obtainable through ultrasonic measurement techniques was performed in this task. The information obtained was then used to determine the uniformity of manufacturing methods and/or the ability to determine non-uniformity in processes.

以某固体火箭发动机推力向量控制系统摆动接头的单珠承载试验模型为计算模型,采用摩擦接触问题的Lagrange乘子法与弹塑性耦合的有限元理论,计算分析了摆动接头阳球试件在不同强化层厚度下的接触应力、变形及破坏机理；为降低系统摆动力矩,同时考虑大尺寸球面的表面强化和加工工艺,提出了满足接触性能的阳球表面强化层为1～1.4mm的厚度要求.通过与单珠承载试验对比分析,检验了有限元建模及算法的合理性；确定了小变形条件下摆动接头弹塑性摩擦接触应力限额为5 384.9 MPa.%In this paper, a single ball load-bearing experiment model of the swing joint of thrust vector control system in a solidpropellant rocket motor was used as computation model. By using the Lagrangian multiplier method for friction contact question, and elasto-plastic couplings finite element method, the contact stress, deformation and destruct principle under different thickness of strengthening layer of the test swing joint convex sphere were computed and analyzed. To reduce system swing moment, and consider manufacturing techniques of large size intensify sphere, 1-1.4 mm thickness of strengthening layer satisfied for contact performance was presented. Rationality of FEM modeling and algorithm was validated by single ball load-bearing experiment. The elasto-plastic friction contact stress of the swing joint was limited to 5384.9 Mpa in small deformation condition.

other non-destructive evaluation ( NDE ) methods is a routine part of many motor programs. Those evaluations do not measure propellant mechanical...6 , was used in the current effort. The model combines the Wilcox k−ω model in the near-wall region with the k− model in the outer part of the...flaw geometry to a far greater extent, since NDE equipment generally do not have high resolution. Therefore, several sensitivity analyses were

Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solidpropellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

Full Text Available Solid rocket motors are propulsion devices for both satellite launchers and missiles, which require guidance and steering to fly along a programmed trajectory and to compensate for flight disturbances. A typical solid rocket motor consists of motor case, solidpropellant grain, motor insulation, igniter and nozzle. In most solid rocket motors, thrust vector control (TVC is required. One of the most efficient methods of TVC is by flex nozzle system. The flex nozzle consists of a flexible bearing made of an elastomeric material alternating with reinforcement rings of metallic or composite material. The material characterisation of AFNOR 15CDV6 steel and the natural rubber-based elastomer developed for use in flex nozzle are discussed. This includes testing, modelling of the material, selection of a material model suitable for analysis, and the validation of material model.Defence Science Journal, 2011, 61(3, pp.264-269, DOI:http://dx.doi.org/10.14429/dsj.61.52

The unstable-resonator spatially enhanced detection (USED) coherent anti-stokes Raman spectroscopy (CARS) measurements of temperature and N2 concentration in the combustion of solidpropellant at atmosphere pressure are reported. The USED CARS measurement system has a high spatial solution of ～ 0.1 mm in diameter and 3 mm in length, and permits instantaneous measurement at 10-Hz rate. The single-pulse N2 Q-branch CARS spectra have been obtained from the propellant combustion. The temperatures and N2 concentrations of the propellant flame at different height have been achieved by fitting the experimental data to theoretical spectra. The results indicate that the temperature is up to～2500 K with N2 concentration in a range from 10% to 26%.

Based on interface properties of HTPB propellant, the cross linking level is regulated through the temperature and time of cure reaction. The chemical reaction takes place gradually in remainder functional groups of the system to form chemical bonds and hydrogen bond, so as to improve the mechanical properties of the resultant. Pre-cure technique and bonding model are described in the paper. They can be used in propellant-liner bonding, propellant loading and integral repairing of the propellant grain of the solid rocket motor (SRM). These test results have been qualified by the successful static firing tests, flight tests of motors and storage tests of specimens for ten-years. The performances of the motor meet the design requirements with good reproducibilities.

Highlights: • Al-Li alloy propellant has increased ideal specific impulse over neat aluminum. • Al-Li alloy propellant has a near complete reduction in HCl acid formation. • Reduction in HCl was verified with wet bomb experiments and DSC/TGA-MS/FTIR. - Abstract: Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solidpropellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (I{sub SP}). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal I{sub SP} by ∼7 s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solidpropellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5 ± 4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption.

A microwave Doppler shift system, with increased resolution over earlier microwave techniques, was developed for the purpose of measuring the regression rates of solidpropellants during rapid pressure transients. A continuous microwave beam is transmitted to the base of a burning propellant sample cast in a metal waveguide tube. A portion of the wave is reflected from the regressing propellant-flame zone interface. The phase angle difference between the incident and reflected signals and its time differential are continuously measured using a high resolution microwave network analyzer and related instrumentation. The apparent propellant regression rate is directly proportional to this latter differential measurement. Experiments were conducted to verify the (1) spatial and time resolution of the system, (2) effect of propellant surface irregularities and compressibility on the measurements, and (3) accuracy of the system for quasi-steady-state regression rate measurements. The microwave system was also used in two different transient combustion experiments: in a rapid depressurization bomb, and in the high-frequency acoustic pressure environment of a T-burner.

The CAEPE, the French Centre of Propellers and Engines Completion and Testing is in charge of the static tests of solidpropellant fuelled propellers. In order to evaluate the risks linked with the explosion of such propellers, an analytical method is proposed which allows to determine the energy, impulses and overpressure due to an explosion occurring at any time of the firing process. This method is based on the exploitation of the shock tube equations. Concerning the overpressure wave propagation, several laws are compared to full scale tests performed at the CAEPE. (J.S.) 14 refs.

Instantaneous burning rate data for a polybutadiene acrylic acid propellant, containing 16 weight percent aluminum, were calculated from the pressure histories of a test motor with 96.77 sq cm of burning area and a 5.08-cm-thick propellant web. Additional acceleration tests were conducted with reduced propellant web thicknesses of 3.81, 2.54, and 1.27 cm. The metallic residue collected from the various web thickness tests was characterized by weight and shape and correlated with the instantaneous burning rate measurements. Rapid depressurization extinction tests were conducted in order that surface pitting characteristics due to localized increased burning rate could be correlated with the residue analysis and the instantaneous burning rate data. The acceleration-induced burning rate augmentation was strongly dependent on propellant distance burned, or burning time, and thus was transient in nature. The results from the extinction tests and the residue analyses indicate that the transient rate augmentation was highly dependent on local enhancement of the combustion zone heat feedback to the surface by the growth of molten residue particles on or just above the burning surface. The size, shape, and number density of molten residue particles, rather than the total residue weight, determined the acceleration-induced burning rate augmentation.

The ablation in solid-propellant rocket nozzle is a coupling process resulted by chemistry, heat and mass transfer. Based on the heat and mass transfer theory, the aero-thermo-dynamic, and thermo-chemical kinetics, the thermal-chemical ablation model is established. Simulations are completed on the heat flow field and chemical ablation in the nozzle with different concentrations, frequency factors and activation energy of H2. The calculation results show that the concentration and the activation energy of H2 can provoke the transformation of control mechanism, whereas the influence brought by the frequency factor of H2 is feeble under a high-temperature and high-pressure combustion circumstance. The discrimination for ablative control mechanism is dependent on both concentration and activation energy of H2. This study will be useful in handling ablation and thermal protection problem in the design of solid-propellant rocket.

A quantitative theory is outlined which allows calculation of crosslink density of solidpropellant binders from a small number of predetermined parameters such as the binder composition, the functionality distributions of the ingredients, and the extent of the curing reaction. The parameter which is partly dependent on process conditions is the extent of reaction. The proposed theoretical model is verified by independent measurement of effective chain concentration and sol and gel fractions in simple compositions prepared from model compounds. The model is shown to correlate tensile data with composition in the case of urethane-cured polyether and certain solidpropellants. A formula for the branching coefficient is provided according to which if one knows the functionality distributions of the ingredients and the corresponding equivalent weights and can measure or predict the extent of reaction, he can calculate the branching coefficient of such a system for any desired composition.

Full Text Available In order to obtain the burn rate of the solidpropellant that is the important parameter of transient burning, the new method named digital image processing is presented. In the article , the principle of digital image processing is analysed; The burning face of the sample in the each time is located according the image and the coordinates of the burning face is obtained. In experiment the transient burn rate is measured by digital image processing and the accuracy is acceptable.

Existing and proposed methods for the prediction of plume radiation are discussed in terms of their application to the NASA Advanced Solid Rocket Motor (ASRM) and Space Shuttle Main Engine (SSME) projects. Extrapolations of the Solid Rocket Motor (SRM) are discussed with respect to preliminary predictions of the primary and secondary radiation environments. The methodology for radiation and initial plume property predictions are set forth, including a new code for scattering media and independent secondary source models based on flight data. The Monte Carlo code employs a reverse-evaluation approach which traces rays back to their point of absorption in the plume. The SRM sea-level plume model is modified to account for the increased radiation in the ASRM plume due to the ASRM's propellant chemistry. The ASRM cycle-1 environment predictions are shown to identify a potential reason for the shutdown spike identified with pre-SRM staging.

Aluminium is widely used as an important additive to improve ballistic and energy performance in solidpropellants, but the unburned aluminium does not contribute to the specific impulse and has both thermal and momentum two-phase flow losses. So understanding of aluminium combustion behaviour during solidpropellant burning is significant when improving internal ballistic performance. Recent developments and experimental results reported on such combustion behaviour are presented in this paper. A variety of experimental techniques ranging from quenching and dynamic measurement, to high-speed CCD video recording, were used to study aluminium combustion behaviour and the size distribution of the initial agglomerates. This experimental investigation also provides the size distribution of the condensed phase products. Results suggest that the addition of an organic fluoride compound to solidpropellant will generate smaller diameter condensed phase products due to sublimation of AlF3. Lastly, a physico-chemical picture of the agglomeration process was also developed based on the results of high-speed CCD video analysis.

The infrared spectral characteristics of high-intensity IR solidpropellant were measured in this paper using a Bruker EQUINOX55 remote sensing FTIR spectrometer. The emission spectra of the combustion flame were recorded in the range of 4700-1800 cm(-1) with a spectral resolution of 4 cm(-1). The combustion temperatures of solidpropellant at the burning time of 0, 9, 18, 27 and 36 s calculated from the molecular rotation-vibration spectra were 1992.5, 1610.9, 2294.4, 2361.1 and 1916.9 K, respectively. Moreover, the spectral radiance distributions of the high IR flare material at different times were given, and the combustion product concentrations of HCl, HF, CO2 and CO were determined quantitatively. Results showed that remote sensing FTIR is a potential technology that can be applied to the measurement of IR spectral characteristics, especially to the identification of the IR objects, guidance and anti-guidance in the military, and the modification of the make-up of solidpropellant.

Full Text Available In rocketry application, now-a-days instead of monopropellants slowly composite propellants are introduced. Burning rate of a solid state composite propellant depends on many factors like oxidizer-binder ratio, oxidizer particle size and distribution, particle size and its distribution, pressure, temperature, etc. Several researchers had taken the mass varied composite propellant. In that, the ammonium perchlorate mainly varied from 85 to 90%. This paper deals with the oxidizer rich propellant by allowing small variation of fuel cum binder ranging from 2%, 4%, 6%, and 8% by mass. Since the percent of the binder is very less compared to the oxidizer, the mixture remains in a powder form. The powder samples are used to make a pressed pellet. Experiments were conducted in closed window bomb set-up at pressures of 2, 3.5, and 7 MN/m2. The burning rates are calculated from the combustion photography (images taken by a high-speed camera. These images were processed frame by frame in MATLAB, detecting the edges in the images of the frames. The burning rate is obtained as the slope of the linear fit from MATLAB and observed that the burn rate increases with the mass variation of constituents present in solid state composite propellant. The result indicates a remarkable increase in burn rate of 26.66%, 20%, 16.66%, and 3.33% for Mix 1, 2, 3, 4 compared with Mix 5 at 7 MN/m2. The percentage variations in burn rate between Mix 1 and Mix 5 at 2, 3.5, and 7 MN/m2 are 25.833%, 32.322%, and 26.185%, respectively.

Commercially available software packages today allow users to quickly perform the routine evaluations of (1) descriptive statistics to numerically and graphically summarize both sample and population data, (2) inferential statistics that draws conclusions about a given population from samples taken of it, (3) probability determinations that can be used to generate estimates of reliability allowables, and finally (4) the setup of designed experiments and analysis of their data to identify significant material and process characteristics for application in both product manufacturing and performance enhancement. This paper presents examples of analysis and experimental design work that has been conducted using Statgraphics®(Registered Trademark) statistical software to obtain useful information with regard to solid rocket motorpropellants and internal insulation material. Data were obtained from a number of programs (Shuttle, Constellation, and Space Launch System) and sources that include solidpropellant burn rate strands, tensile specimens, sub-scale test motors, full-scale operational motors, rubber insulation specimens, and sub-scale rubber insulation analog samples. Besides facilitating the experimental design process to yield meaningful results, statistical software has demonstrated its ability to quickly perform complex data analyses and yield significant findings that might otherwise have gone unnoticed. One caveat to these successes is that useful results not only derive from the inherent power of the software package, but also from the skill and understanding of the data analyst.

The motor case used in the solidpropellant booster for the Space Shuttle is unique in many respects, most of which are indigenous to size and special design requirements. The evolution of the case design from initial requirements to finished product is discussed, with increased emphasis of reuse capability, special design features, fracture mechanics and corrosion control. Case fabrication history and the resulting procedure are briefly reviewed with respect to material development, processing techniques and special problem areas. Case assembly, behavior and performance during the DM-1 static firing are reviewed, with appropriate comments and conclusions.

Hundred-meter-sized objects have been identified by the Cassini spacecraft in Saturn's A ring through the so-called propeller features they create in the ring. These moonlets should migrate due to their gravitational interaction with the ring; in fact, some orbital variations have been detected. The standard theory of type I migration of planets in protoplanetary disks cannot be applied to the ring system as it is pressureless. Thus, we compute the differential torque felt by a moonlet embedded in a two-dimensional disk of solid particles, with a flat surface density profile, both analytically and numerically. We find that the corresponding migration rate is too small to explain the observed variations of the propeller's orbit in Saturn's A ring. However, local density fluctuations (due to gravity wakes in the marginally gravitationally stable A ring) may exert a stochastic torque on a moonlet. Our simulations show that this torque can be large enough to account for the observations depending on the parameters of the rings. We find that on timescales of several years the migration of propellers is likely to be dominated by stochastic effects (while the former, non-stochastic migration dominates after ~104-105 years). In that case, the migration rates provided by observations so far suggest that the surface density of the A ring should be on the order of 700 kg m-2. The age of the propellers should not exceed 1-100 million years depending on the dominant migration regime.

Full Text Available Self-organization phenomena are of critical importance in living organisms and of great interest to exploit in nanotechnology. Here we describe in vitro self-organization of molecular motor-propelled actin filaments, manifested as a tendency of the filaments to accumulate in high density close to topographically defined edges on nano- and microstructured surfaces. We hypothesized that this “edge-tracing” effect either (1 results from increased motor density along the guiding edges or (2 is a direct consequence of the asymmetric constraints on stochastic changes in filament sliding direction imposed by the edges. The latter hypothesis is well captured by a model explicitly defining the constraints of motility on structured surfaces in combination with Monte-Carlo simulations [cf. Nitta et al. (2006] of filament sliding. In support of hypothesis 2 we found that the model reproduced the edge tracing effect without the need to assume increased motor density at the edges. We then used model simulations to elucidate mechanistic details. The results are discussed in relation to nanotechnological applications and future experiments to test model predictions.

Traditionally, Solid Rocket Motor (SRM) internal ballistic performance has been analyzed and predicted with either zero-dimensional (volume filling) codes or one-dimensional ballistics codes. One dimensional simulation of SRM performance is only necessary for ignition modeling, or for motors that have large length to port diameter ratios which exhibit an axial "pressure drop" during the early burn times. This type of prediction works quite well for many types of motors, however, when motor aspect ratios get large, and port to throat ratios get closer to one, two dimensional effects can become significant. The initial propellant grain configuration for the Space Shuttle Reusable Solid Rocket Motor (RSRM) was analyzed with 2-D, steady, axi-symmetric computational fluid dynamics (CFD). The results of the CFD analysis show that the steady-state performance prediction at the initial burn geometry, in general, agrees well with 1-D transient prediction results at an early time, however, significant features of the 2-D flow are captured with the CFD results that would otherwise go unnoticed. Capturing these subtle differences gives a greater confidence to modeling accuracy, and additional insight with which to model secondary internal flow effects like erosive burning. Detailed analysis of the 2-D flowfield has led to the discovery of its hidden 1-D isentropic behavior, and provided the means for a thorough and simplified understanding of internal solid rocket motor flow. Performance parameters such as nozzle stagnation pressure, static pressure drop, characteristic velocity, thrust and specific impulse are discussed in detail and compared for different modeling and prediction methods. The predicted performance using both the 1-D codes and the CFD results are compared with measured data obtained from static tests of the RSRM. The differences and limitations of predictions using ID and 2-D flow fields are discussed and some suggestions for the design of large L/D motors and

Full Text Available This study was designed to evaluate the effect of propellant formulation and geometry on the solidpropellant grains internal ballistic performance using core, bates, rod and tubular and end-burn geometries. Response Surface Methodology (RSM was used to analyze and optimize the effect of sucrose, potassium nitrate and carbon on the chamber pressure, temperature, thrust and specific impulse of the solidpropellant grains through Central Composite Design (CCD of the experiment. An increase in potassium nitrate increased the specific impulse while an increase in sucrose and carbon decreased specific impulse. The coefficient of determination (R2 for models of chamber pressure, temperature, thrust and specific impulse in terms of composition and geometry were 0.9737, 0.9984, 0.9745 and 0.9589, respectively. The optimum specific impulse of 127.89 s, pressure (462201 Pa, temperature (1618.3 K and thrust (834.83 N were obtained using 0.584 kg of sucrose, 1.364 kg of potassium nitrate and 0.052 kg of carbon as well as bate geometry. There was no significant difference between the calculated and experimented ballistic properties at p < 0.05. The bate grain geometry is more efficient for minimizing the oscillatory pressure in the combustion chamber.

A theory for combustion of solidpropellants in a closed vessel is presented allowing for residual exothermic chemical reaction in the bulk of the gas in the vessel. Particular attention is given to propellants exhibiting thick gaseous flame zones such as nitrocellulose, double-base and nitramine propellants. For these, the reaction at high pressures is assumed to involve mainly the oxidation of residual hydrocarbons by NO. It is shown that the direct dynamic coupling between the exothermicity, the molecular weight reduction and the changing pressure can influence the dp/dt-p traces obtained, in a manner not directly related to mass burning rate of the solid. Energy and species conservation equations are derived for the bulk of the vessel in differential form; the system is solved numerically. The results show the effect of extended chemical reaction upon measurable combustion characteristics such as dp/dt-p and burn rate pressure exponent, demonstrating its potential importance in interpretation of closed vessel firing data, depending on the pace of the residual gas phase reactions.

By using fractal geometry is is possible to calculate the actual AP(Ammonium Perchlorate)surface area and oxidizer-binder interface fractal dimension in the prediction of burning rate of commposite solidpropellants.In this investigation,the fractal dimension was determined by a procedure known as the "Box counting Method".using this dimensio,surface area relations were developed for the rough particles.This method was implemented in the PEM(Petite Ensemble Model) burning rate model,The comparison of burning rates for a typical propellant by the PEM and fractal model shows that the burning rates botained by using the fractal geometry are slightly less than those obtained by the PEM model.

A mathematical model is developed to predict burn rate augmentation due to strain in a composite solidpropellant. The model assumes the effect is due to the ability of the flame to penetrate the small fissures and voids that form when a propellant is strained. The number and size of these fissures is obtained by applying a flaw propagation analysis to randomly distributed flaws that form when the binder-oxidizer particle bonds break under stress. A flame height is calculated with Summerfield's burn rate equation and is used to compute the burn rate augmentation based upon the additional burn area created when the flame penetrates the fissures. Comparisons are made with data obtained from published sources. The existence of threshold pressure and strains, above which augmentation occurs, is verified although the model predicts a lower threshold pressure and higher threshold strain than expected. Further results and applications of the model are discussed.

This paper presents particle formation energy balances and detailed analyses of the images from experiments that were conducted on the formation of solid hydrogen particles in liquid helium during the Phase II testing in 2001. Solid particles of hydrogen were frozen in liquid helium and observed with a video camera. The solid hydrogen particle sizes and the total mass of hydrogen particles were estimated. The particle formation efficiency is also estimated. Particle sizes from the Phase I testing in 1999 and the Phase II testing in 2001 were similar. Though the 2001 testing created similar particles sizes, many new particle formation phenomena were observed. These experiment image analyses are one of the first steps toward visually characterizing these particles and it allows designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Full Text Available The ability to predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave development in the motor chamber flow. With the focus of the present study placed on reactive particles, a numerical internal ballistic model incorporating relevant elements, such as a transient, frequency-dependent combustion response to axial pressure wave activity above the burning propellant surface, is applied to the investigation of using aluminum particles within the central internal flow (particles whose surfaces nominally regress with time, as a function of current particle size, as they move downstream as a means of suppressing instability-related symptoms in a cylindrical-grain motor. The results of this investigation reveal that the loading percentage and starting size of the aluminum particles have a significant influence on reducing the resulting transient pressure wave magnitude.

It was Zel'dovich, who first considered the transient combustion problem of solidpropellants. Some more detailed models of that process have been developed afterwards. However, until today, numerical methods remain the prevailing tool for modeling unsteady combustion processes. In this work, it has been demonstrated that at least one of the problems of the unsteady combustion theory, which previously investigated numerically, can be treated analytically by means of fractional calculus. The solution for the unsteady speed of combustion thus derived is then compared with the solution obtained by numerical means in previous studies. The comparison shows a good agreement between those results, especially for small values of time.

Polymeric binders for solidpropellants are usually based on hydroxyl-terminated polybutadiene (HTPB), which does not contribute to the overall energy output. Azidic polyethers represent an interesting alternative but may have poorer mechanical properties. Polybutadiene-polyether copolymers may combine the advantages of both. Four different ether-butadiene-ether triblock copolymers were prepared and azidated starting from halogenated and/or tosylated monomers using HTPB as initiator. The presence of the butadiene block complicates the azidation step and reduces the storage stability of the azidic polymer. Nevertheless, the procedure allows modifying the binder properties by varying the type and lengths of the energetic blocks.

The Reusable Solid Rocket Motor represents the largest solid rocket motor ever flown and the only human rated solidmotor. Each Reusable Solid Rocket Motor (RSRM) provides approximately 3-million lb of thrust to lift the integrated Space Shuttle vehicle from the launch pad. The motors burn out approximately 2 minutes later, separate from the vehicle and are recovered and refurbished. The size of the motor and the need for high reliability were challenges. Thrust shaping, via shaping of the propellant grain, was needed to limit structural loads during ascent. The motor design evolved through several block upgrades to increase performance and to increase safety and reliability. A major redesign occurred after STS-51L with the Redesigned Solid Rocket Motor. Significant improvements in the joint sealing systems were added. Design improvements continued throughout the Program via block changes with a number of innovations including development of low temperature o-ring materials and incorporation of a unique carbon fiber rope thermal barrier material. Recovery of the motors and post flight inspection improved understanding of hardware performance, and led to key design improvements. Because of the multidecade program duration material obsolescence was addressed, and requalification of materials and vendors was sometimes needed. Thermal protection systems and ablatives were used to protect the motor cases and nozzle structures. Significant understanding of design and manufacturing features of the ablatives was developed during the program resulting in optimization of design features and processing parameters. The project advanced technology in eliminating ozone-depleting materials in manufacturing processes and the development of an asbestos-free case insulation. Manufacturing processes for the large motor components were unique and safety in the manufacturing environment was a special concern. Transportation and handling approaches were also needed for the large

The results of a re-evaluation of the propellant combustion data obtained using the dual valve approach for measuring velocity-coupling characteristics of solidpropellants are presented. Data analysis and testing procedures are described. The velocity response is compared to pressure-coupled response data within the context of thermal wave response theory. This comparison shows important inconsistencies which cast doubt on inferring the velocity response from pressure-coupled response functions.

Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor

Full Text Available The calorimetric value (cal-val of solid rocket propellants and explosives is determined in the presence of inert atmosphere using industrial nitrogen gas. However, due to presence of trace amount of oxygen, the cal-val is not always correct. To avoid such inaccuracy in cal-val, a systematic study has been carried out by taking different types of solidpropellant samples having burning rate in the range of 5 mm/s − 30 mm/s at different pressures. The data obtained were acquired using industrial nitrogen, ultra high pure nitrogen (UHP-N2, ultra high pure argon (UHP-Ar, air and ultra high pure oxygen (UHP-O2. The data reveal that cal-val is highest in the case of UHP-O2 due to complete combustion while in the case of air and industrial nitrogen it is found to be substantially less. Moreover, the cal-val in the presence of UHP-N2 and UHP-Ar meets the standard value with reproducibility. The results, further, confirm that for authentic cal-val, the most suitable environment is UHP-N2/UHP-Ar.

Full Text Available The calorimetric value (cal-val of solid rocket propellants and explosives is determined in the presenceof inert atmosphere using industrial nitrogen gas. However, due to presence of trace amount of oxygen, the cal-val is not always correct. To avoid such inaccuracy in cal-val, a systematic study has been carried out by takingdifferent types of solidpropellant samples having burning rate in the range of 5 mm/s − 30 mm/s at different pressures.The data obtained were acquired using industrial nitrogen, ultra high pure nitrogen (UHP-N2, ultra high pureargon (UHP-Ar, air and ultra high pure oxygen (UHP-O2. The data reveal that cal-val is highest in the case ofUHP-O2 due to complete combustion while in the case of air and industrial nitrogen it is found to be substantiallyless. Moreover, the cal-val in the presence of UHP-N2 and UHP-Ar meets the standard value with reproducibility.The results, further, confirm that for authentic cal-val, the most suitable environment is UHP-N2/UHP-Ar.Defence Science Journal, 2013, 63(5, pp.467-472, DOI:http://dx.doi.org/10.14429/dsj.63.2896

This report presents particle formation observations and detailed analyses of the images from experiments that were conducted on the formation of solid hydrogen particles in liquid helium. Hydrogen was frozen into particles in liquid helium, and observed with a video camera. The solid hydrogen particle sizes and the total mass of hydrogen particles were estimated. These newly analyzed data are from the test series held on February 28, 2001. Particle sizes from previous testing in 1999 and the testing in 2001 were similar. Though the 2001 testing created similar particles sizes, many new particle formation phenomena were observed: microparticles and delayed particle formation. These experiment image analyses are some of the first steps toward visually characterizing these particles, and they allow designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Full Text Available Polyurethane propellants, which constitute one of the 'work-horse' binder systems in modern solid rocketry are easily amenable for tailoring the mechanical properties in terms of variations in the molecular structure of the backbone polyols, the isocyanates and stoichiometry of the reactants. The paper deals with studies in developing an advanced binder system based on poly (oxy propylene glycol and toluene-di-isocyanate, which is capable of accommodating high solids loading and conceding elongatlon at maximum stress, of more than 125 per cent. The gum-stock properties of the binder are related with those of a low molecular weight version and the results are explained based on the network theory of condensation polymers.

Proton NMR images of solidpropellant materials, consisting of a polybutadiene binder material filled with 82% solid particles, have been obtained at a magnetic field strength of 14.1 T and at a resolution of 8.5 x 8.5 micron. The images are the first of elastomeric materials obtained at a proton frequency of 600 MHz and have the highest spatial resolution yet reported. The images display a high contrast and are rich in information content. They reveal the distribution of individual filler particles in the polymer matrix as well as a thin polymer film of about 10-30 micron which is found to surround some of the larger filler particles.

The knowledge of the aluminum oxide particle size distribution inside the combustion chamber of a solidpropellant rocket motor is an important factor for assessing the combustion stability or the slag mass accumulation in the motor. A representative subscale motor for the Ariane 5 P230 Solid Rocket Booster (SRB), in which helium is injected to quench the condensed phase reactions, has been designed and manufactured. Its use for combustion stability purpose has given the aluminum oxide particle size distribution in conditions representative of the actual Ariane 5 SRB. The experimental techniques, optical and particle capturing, have been found to give results in good agreement. A stretched distribution, with particles ranging from 1 micron to 120 microns and a maximum around 45 microns, has been demonstrated.

The combinatorial nature of many important mathematical problems, including nondeterministic-polynomial-time (NP)-complete problems, places a severe limitation on the problem size that can be solved with conventional, sequentially operating electronic computers. There have been significant efforts in conceiving parallel-computation approaches in the past, for example: DNA computation, quantum computation, and microfluidics-based computation. However, these approaches have not proven, so far, to be scalable and practical from a fabrication and operational perspective. Here, we report the foundations of an alternative parallel-computation system in which a given combinatorial problem is encoded into a graphical, modular network that is embedded in a nanofabricated planar device. Exploring the network in a parallel fashion using a large number of independent, molecular-motor-propelled agents then solves the mathematical problem. This approach uses orders of magnitude less energy than conventional computers, thus addressing issues related to power consumption and heat dissipation. We provide a proof-of-concept demonstration of such a device by solving, in a parallel fashion, the small instance {2, 5, 9} of the subset sum problem, which is a benchmark NP-complete problem. Finally, we discuss the technical advances necessary to make our system scalable with presently available technology.

The development of the Vulcain program was ensured thanks to a European cooperation with an ESA (European Space Agency) financing. The CNES (European Centre for Space Studies) has ensured the technical and financial direction of the program and gave the control of the development to the SEP. The manufacturing of the Vulcain engine is managed under the Arianespace contract, in charge of the marketing of the Ariane 5 launcher. The overall engineering of the engine and its tests were carried out by the SEP in Vernon (France) and Lampoldshausen (Germany) test facilities. SEP has also developed and built the hydrogen turbo-pump, the gas generator and its feeding valves. Several companies are involved in the development of this engine: DASA for the combustion chamber, Fiat Avio for the oxygen turbo-pump, Volvo Aero Corp. for the divergent and the hydrogen and oxygen turbines, Techspace Aero for the chamber injection valves and the drain and hot gases valves, Microtechnica for the electro-valves and check valves, SPE for the firing and start-up equipments, Avica for the feeding lines, Devtec for the supports, and MAN for the cardan and the thermal protection. This paper describes the functioning principle of the Vulcain engine, and of the two solidpropellant fuel acceleration stages of the Ariane 5 launcher. Some future projects of the SEP are also described: the dual liquid-fuels engine and the plasma engine. (J.S.)

Saturethane solidpropellant was used in all tests. The spore inoculum was evenly distributed in the propellant. Samples weighing approximately 5 g were aseptically removed, placed into curing ovens, and exposed to cure temperatures. Initial tests were conducted at 82 and 93 C. Analysis of the obtained data indicated that the survivor curves were not linear. Exposure of the inoculated propellant samples to 93 C reduced the initial population to less than 0.01% in about 20 hours. At 82 C, approximately 168 hours were required for a similar reduction. Tests involving curing temperatures of 105 and 115 C were also conducted. It is pointed out that changes in the mechanism of spore inactivation due to chemical and physical changes in the propellant could account for the nonlinear survivor curves obtained.

Hundred meter sized objects have been identified by the Cassini spacecraft in Saturn's A ring through the so-called "propeller" features they create in the ring. These moonlets should migrate, due to their gravitational interaction with the ring ; in fact, some orbital variation have been detected. The standard theory of type I migration of planets in protoplanetary disks can't be applied to the ring system, as it is pressureless. Thus, we compute the differential torque felt by a moonlet embedded in a two-dimensional disk of solid particles, with flat surface density profile, both analytically and numerically. We find that the corresponding migration rate is too small to explain the observed variations of the propeller's orbit in Saturn's A-ring. However, local density fluctuations (due to gravity wakes in the marginally gravitationally stable A-ring) may exert a stochastic torque on a moonlet. Our simulations show that this torque can be large enough to account for the observations, depending on the paramet...

A multi-dimensional numerical model has been developed for the unsteady state oscillatory combustion of solidpropellants subject to acoustic pressure disturbances. Including the gas phase unsteady effects, the assumption of uniform pressure across the flame zone, which has been conventionally used, is relaxed so that a higher frequency response in the long flame of a double-base propellant can be calculated. The formulation is based on a premixed, laminar flame with a one-step overall chemical reaction and the Arrhenius law of decomposition with no condensed phase reaction. In a given geometry, the Galerkin finite element solution shows the strong resonance and damping effect at the lower frequencies, similar to the result of Denison and Baum. Extended studies deal with the higher frequency region where the pressure varies in the flame thickness. The nonlinear system behavior is investigated by carrying out the second order expansion in wave amplitude when the acoustic pressure oscillations are finite in amplitude. Offset in the burning rate shows a negative sign in the whole frequency region considered, and it verifies the experimental results of Price. Finally, the velocity coupling in the two-dimensional model is discussed.

Full Text Available The aim of this work was to determine the dioctyl azelate (DOZ plasticizer diffusion coefficient (D for samples containing the interfaces of rubber, liner and solid composite propellant based on hydroxyl-terminated polybutadiene (HTPB. The samples used in the diffusion study were aged up to 31 days after the cure period at 80 °C. A computer program implementing a mathematical model of Fick's second Law of diffusion was developed to calculate the diffusion coefficient based on concentration data obtained from gas chromatographic analyses. The effects of the diffusion phenomenon were also investigated by Shore A hardness and scanning electron microscope (SEM techniques. These analyses were carried out using samples aged at room temperature and at 80 °C. The hardness results showed an increasing trend for the samples aged at room temperature; however in the tests carried out at 80 °C they showed the opposite trend. The SEM analyses detected meaningful changes in the surface morphology of the propellant for both aging temperatures.

In order to study the damage evolution law of composite solidpropellants, the molecular dynamics particle filled algorithm was used to establish the mesoscopic structure model of HTPB(Hydroxyl-terminated polybutadiene) propellants. The cohesive element method was employed for the adhesion interface between AP(Ammonium perchlorate) particle and HTPB matrix and the bilinear cohesive zone model was used to describe the mechanical response of the interface elements. The inversion analysis method based on Hooke-Jeeves optimization algorithm was employed to identify the parameters of cohesive zone model(CZM) of the particle/binder interface. Then, the optimized parameters were applied to the commercial finite element software ABAQUS to simulate the damage evolution process for AP particle and HTPB matrix, including the initiation, development, gathering and macroscopic crack. Finally, the stress-strain simulation curve was compared with the experiment curves. The result shows that the bilinear cohesive zone model can accurately describe the debonding and fracture process between the AP particles and HTPB matrix under the uniaxial tension loading.

This paper describes our work to magnetically steer self-propelled devices whichmove by catalysis of hydrogen peroxide. We demonstrate manipulation of paramagnetic particles with a diameter of 100μm to a given setpoint by means of magnetic field gradients, as well as self-propelled movement of gold-

提出了固体碳氢推进剂作为涡轮增压固冲发动机( TSPR)驱涡推进剂的方案,分析了适用于TSPR推进剂的热力参数和一次燃烧产物成分,完成了驱涡推进剂的选择;进行了备选推进剂( CH04)对TSPR性能的影响性评估,证明该推进剂能够满足TSPR的性能要求;对所选推进剂了进行了一、二次燃烧试验,试验结果表明,CH04推进剂在补燃室点火较困难,但其一次、二次燃烧稳定性好,燃气参数基本满足TSPR对推进剂性能要求.%Solid hydrocarbon propellant was proposed to be used to drive turbine of Turbocharged SolidPropellant Ramjet (TSPR).The thermodynamic parameters and the first combustion production of propellant used for TSPR were analyzed to select the suitable propellant(CH04) to drive the turbine. An evaluation for propellant' s influence on the performance of TSPR was per-formed,which proved that the propellant(CH04)meets the requirements of TSPR.The first and second combustion experiments were carried out.The results show that,the first and second combustion have high stability,despite of the difficulty in ignition in the after-burner,and the feasibility of the selected propellant used for TSPR was verified.

A reasonable discrimination method for ablative control mechanism in solid-propellant rocket nozzle can improve the calculation accuracy of ablation rate. Based on the different rate constants for reactions of C with H2O and CO2,a new discrimination method for ablative control mechanism,which comprehensively considers the influence of nozzle surface temperature and gas component concentration,is presented. Using this new discrimination method,calculations were performed to simulate the nozzle throat insert ablation. The numerical results showed that the calculated ablation rate,which was more close to the measured values,was less than the value calculated by diffusion control mechanisms or by double control mechanisms. And H2O was proved to be the most detrimental oxidizing species in nozzle ablation.

The investigation aims at the expansion of the basis of formulations of solid composite propellants by introducing new compositions with lower sensitivity to mechanic impact and improved thermal stability .The formulations based on trinitropyrazole (TNP) contains a binder (a hydrocarbon or active one) ,aluminum and inorganic oxidizer ADN .The results show that a binary for‐mulation TNP + active binder (18% -19% )(volume fraction) with no metal is well designed which would achieve high specific im‐pulse (at Pc∶ Pa=40∶1) of 248 s ,high density of 1 .80 g/cm3 and combustion temperature Tc about 3450 K .In terms of energy , metal‐free compositions with TNP lose a bit to those with HMX ,only if HMX fraction in formulation is higher than 45% -50% .

This paper presents a functional on-chip pressure generator that utilizes chemical energy from a solid chemical propellant to perform fluidic delivery in applications of plastic-based disposable biochips or lab-on-a-chip systems. In this functional on-chip pressure generator, azobis-isobutyronitrile (AIBN) as the solid chemical propellant is deposited on a microheater using a screen-printing technique, which can heat the AIBN at 70 degrees C to produce nitrogen gas. The output pressure of nitrogen gas, generated from the solid chemical propellant, is adjustable to a desired pressure by controlling the input power of the heater. Using this chemical energy source, the generated pressure depends on the deposited amount of the solid chemical propellant and the temperature of the microheater. Experimental measurements show that this functional on-chip pressure generator can achieve around 3 000 Pa pressure when 189 mJ of energy is applied to heat the 100 microg of AIBN. This pressure can drive 50 nl of water through a microfluidic channel of 70 mm and cross-sectional area of 100 microm x 50 microm. Due to its compact size, ease of fabrication and integration, high reliability (no moving parts), biologically inert gas output along with functionality of gas generation, this pressure generator will be an excellent pressure source for handling the fluids of disposable lab-on-a-chip, biochemical analysis systems or drug delivery systems.

.... The plasma jet velocity 2 inches from the source exit was found to be about 1300 m/s. Following characterization of the plasma-flow field, a set of experiments was conducted on JA-2 solidpropellant with controlled bed temperature...

In a solid rocket motor, nozzle throat erosion is dictated by propellant composition, throat material properties, and operating conditions. Throat erosion has a significant effect on motor performance, so it must be accurately characterized to produce a good motor design. In order to correlate throat erosion rate to other parameters, it is first necessary to know what the throat diameter is throughout a motor burn. Thus, an indirect method and a direct method for determining throat diameter in a solid rocket motor are investigated in this thesis. The indirect method looks at the use of pressure and thrust data to solve for throat diameter as a function of time. The indirect method's proof of concept was shown by the good agreement between the ballistics model and the test data from a static motor firing. The ballistics model was within 10% of all measured and calculated performance parameters (e.g. average pressure, specific impulse, maximum thrust, etc.) for tests with throat erosion and within 6% of all measured and calculated performance parameters for tests without throat erosion. The direct method involves the use of x-rays to directly observe a simulated nozzle throat erode in a dynamic environment; this is achieved with a dynamic calibration standard. An image processing algorithm is developed for extracting the diameter dimensions from the x-ray intensity digital images. Static and dynamic tests were conducted. The measured diameter was compared to the known diameter in the calibration standard. All dynamic test results were within +6% / -7% of the actual diameter. Part of the edge detection method consists of dividing the entire x-ray image by an average pixel value, calculated from a set of pixels in the x-ray image. It was found that the accuracy of the edge detection method depends upon the selection of the average pixel value area and subsequently the average pixel value. An average pixel value sensitivity analysis is presented. Both the indirect

This project is to develop and test a new propellant formulation specifically for the Mars Ascent Vehicle (MAV) for the robotic Mars Sample Return mission. The project was initiated under the Planetary Sciences Division In-Space Propulsion Technology (ISPT) program and is continuing under the Mars Exploration Program. The two-stage, solidmotor-based MAV has been the leading MAV solution for more than a decade. Additional studies show promise for alternative technologies including hybrid and bipropellant options, but the solidmotor design has significant propellant density advantages well suited for physical constraints imposed while using the SkyCrane descent stage. The solidmotor concept has lower specific impulse (Isp) than alternatives, but if the first stage and payload remain sufficiently small, the two-stage solid MAV represents a potential low risk approach to meet the mission needs. As the need date for the MAV slips, opportunities exist to advance technology with high on-ramp potential. The baseline propellant for the MAV is currently the carboxyl terminated polybutadiene (CTPB) based formulation TP-H-3062 due to its advantageous low temperature mechanical properties and flight heritage. However, the flight heritage is limited and outside the environments, the MAV must endure. The ISPT program competed a propellant formulation project with industry and selected ATK to develop a new propellant formulation specifically for the MAV application. Working with ATK, a large number of propellant formulations were assessed to either increase performance of a CTPB propellant or improve the low temperature mechanical properties of a hydroxyl terminated polybutadiene (HTPB) propellant. Both propellants demonstrated potential to increase performance over heritage options, but an HTPB propellant formulation, TP-H-3544, was selected for production and testing. The test plan includes propellant aging first at high vacuum conditions, representative of the Mars transit

The three-dimensional numerical simulation of two-phase plume flow of solidpropellant micro-thrusters was developed.Then it was used to investigate the plume interference effect by combining the direct simulation Monte Carlo（DSMC） method for multi-component gas flow with the two-way coupling model for two-phase rarefied flow.At different space between the two micro-thrusters and different wall temperature,the plume interference effect was analyzed specifically.The results show that under the plume interference effect the gas is compressed and the flow direction is changed,which resulted in the increasing of gas pressure and temperature;solid phase made no significant effect on the flow parameters of gas phase;with the rising of the space between the two micro-thrusters,the maximum pressure decreased and the maximum temperature increased in the domain under the plume interference effect;the wall temperature could influence the temperature of the gas which is extremely close to the wall,but not the gas pressure.

The River Protection Project (RPP) is planning to retrieve radioactive waste from the single-shell tanks (SST) and double-shell tanks (DST) underground at the Hanford Site. This waste will then be transferred to a waste treatment plant to be immobilized (vitrified) in a stable glass form. Over the years, the waste solids in many of the tanks have settled to form a layer of sludge at the bottom. The thickness of the sludge layer varies from tank to tank, from no sludge or a few inches of sludge to about 15 ft of sludge. The purpose of this technology and engineering case study is to evaluate the Flygt{trademark} submersible propeller mixer as a potential technology for auxiliary mobilization of DST HLW solids. Considering the usage and development to date by other sites in the development of this technology, this study also has the objective of expanding the knowledge base of the Flygt{trademark} mixer concept with the broader perspective of Hanford Site tank waste retrieval. More specifically, the objectives of this study delineated from the work plan are described.

Full Text Available This work aims to determine the diffusion coefficient of the plasticizers dibutyl phthalate (DBP, dioctyl phthalate (DOP and dioctyl azelate (DOZ on the internal insulating layer of solid rocket motors. These plasticizers are originally present in the layers of rubber, liner and propellant, respectively. This species are not chemically bonded and tend to diffuse from propellant to insulating and vice versa. A computer program based on the mathematical model of Fick’s second Law of diffusion was developed to perform the calculus from the concentration data obtained by gas chromatographic (GC analyses. The samples were prepared with two different adhesive liners; one conventional (LHNA and the other with barrier properties (LHNT. A common feature of both liners was that they were synthesized by the reaction of hydroxyl-terminated polybutadiene (HTPB and diisocyanates. However, a bond promoter was used to increase the crosslink density of the LHNT liner and to improve its performance as barrier against the diffusion. The effects of the diffusion of the plasticizers were also investigated by hardness analyses, which were executed on samples aged at room temperature and at 80ºC. The results showed an increase trend for the samples aged at room temperature and an opposite behavior for the tests carried out at 80ºC.

The Black Brant variation of the Standard Brant developed in the 1960's has been a workhorse motor of the NASA Sounding Rocket Project Office (SRPO) since the 1970's. In March 2012, the Black Brant Mk1 used on mission 36.277 experienced combustion instability during a flight at White Sands Missile Range, the third event in the last four years, the first occurring in November, 2009, the second in April 2010. After the 2010 event the program has been increasing the motor's throat diameter post-delivery with the goal of lowering the chamber pressure and increasing the margin against combustion instability. During the most recent combustion instability event, the vibrations exceeded the qualification levels for the Flight Termination System. The present study utilizes data generated from T-burner testing of multiple Black Brant propellants at the Naval Air Warfare Center at China Lake, to improve the combustion stability predictions for the Black Brant Mk1 and to generate new predictions for the Mk2. Three unique one dimensional (1-D) stability models were generated, representing distinct Black Brant flights, two of which experienced instabilities. The individual models allowed for comparison of stability characteristics between various nozzle configurations. A long standing "rule of thumb" states that increased stability margin is gained by increasing the throat diameter. In contradiction to this experience based rule, the analysis shows that little or no margin is gained from a larger throat diameter. The present analysis demonstrates competing effects resulting from an increased throat diameter accompanying a large response function. As is expected, more acoustic energy was expelled through the nozzle, but conversely more acoustic energy was generated due to larger gas velocities near the propellant surfaces.

An exact analytical solution is given to the problem of long-time propellant thermal response to a specified pressure oscillation. Coupling to the gas phase is made using the quasisteady Zeldovich-Novozhilov approximation. Explicit linear and lowest order (quadratic) nonlinear expressions for propellant response are obtained from the implicit nonlinear solutions. Using these expressions, response curves are presented for an ammonium perchlorate composite propellant and HMX monopropellant.

Full Text Available Studies on Thermal decomposition of ammonium perchlorarte(AP- polystyrene(PS propellant and burning rate of PS/AP propellant have been carried out as a function of oxidizer particle size. Thermal decomposition of AP and AP/PS propellant as function of AP particle size shows a maximum rate around 100 micro particle size which has been explained on the basis of Mample's theory. No such maximum is observed in the case of PS/AP propellant burning rate.

Igniter is the basic component of MEMS-based solidpropellant microthrusters (SPM) array, which should response fast with low dissipated energy. To satisfy the requirements, micro-semiconductor bridge (MSCB) with dual V-angles contributing to the reduction of dissipated energy is introduced into the array. The electrical explosion characteristics of MSCB investigated under constant voltage, in the considerations of the limit of power supply in the micro/nanosatellite, are similar to those under capacitive discharge. The bridge was completely vaporized and burst into the hot plasma under high-level voltage. While under low-level voltage, the bridge was partly vaporized without detected plasma. The 3#SCB (90° V-angle) with the smallest power density resulting in the lowest applied voltage (12.0 V) and the 2#SCB with the smallest narrow width (Wmin) leading to the lowest average power (11.3 W) were chosen to the further optimization. The ignition tests were implemented to examine the ignition reliability of the contractible igniters, which were optimized to satisfy the voltage limitation with sufficient ignition capacity. The results show that the effective SCB with lowest dissipated power is the 3#-3SCB, which function within 100 μ under 7.0 V, and the average power below 5.0 W.

A viscoelastic deformation and damage model (VED) for solid rocket propellants has been developed based on an extensive set of mechanical properties experiments. Monotonic tensile tests performed at several strain rates showed rate and dilatation effects. During cyclic tensile tests, hysteresis and a rate-dependent shear modulus were observed. A tensile relaxation experiment showed significant stress decay in the sample. Taylor impact tests exhibited large dilatations without significant crack growth. Extensive modifications to a viscoelastic-viscoplastic model (VEP) necessary to capture these experimental results have led to development of the VED model. In particular, plasticity has been eliminated in the model, and the multiple Maxwell viscoelastic formulation has been replaced with a time-dependent shear modulus. Furthermore, the loading and unloading behaviors of the material are modeled independently. To characterize the damage and dilatation behavior, the Tensile Damage and Distention (TDD) model is run in conjunction with VED. The VED model is connected to a single-cell driver as well as to the CTH shock physics code. Simulations of tests show good comparisons with tensile tests and some aspects of the Taylor tests.

Full Text Available The paper deals with various theoretical approaches to the mathematical modeling of the operating process in solidpropellant ramjets (SPRJ that use highly metalized solidpropellant. It introduces a new method (combustion operating law method that allows us to carry out comparative analysis of combustion efficiency in SPRJ arbitrary geometry ram-burners (RB when there is no accurate information on the combustion law of condensed fuel particles. To illustrate an application of the proposed method, mathematical modeling of the operating process was conducted for three SPRJ ram-burners with three different air intakes (AI, for which distribution fields of main parameters of gas and fuel particles have been obtained. Most complete combustion of fuel particles and the lowest level of particles buildup are registered for RB180 (180 degree angle between AIs. The results of a comparative analysis show that the relative (compared to RB180 efficiency of the particle burning process equals 0.64 and 0.6, respectively, for RB90 (90 degree angle between AIs and RB60 (60 degree angle between AIs. The proposed method may be applied to solve the most difficult problems of mathematical modeling when the optimization development of the solidpropellant and ramjet structure are fulfilled simultaneously, i.e. when designers do not have the complete information about the combustion law of the condensed fuel particles.

Space shuttle solid rocket motor case assembly joints are sealed with conventional O-ring seals that are shielded from 5500 F combustion gases by thick layers of insulation and by special joint-fill compounds that fill assembly splitlines in the insulation. On a number of occasions, NASA has observed hot gas penetration through defects in the joint-fill compound of several of the rocket nozzle assembly joints. In the current nozzle-to-case joint, NASA has observed penetration of hot combustion gases through the joint-fill compound to the inboard wiper O-ring in one out of seven motors. Although this condition does not threaten motor safety, evidence of hot gas penetration to the wiper O-ring results in extensive reviews before resuming flight. The solid rocket motor manufacturer (Thiokol) approached the NASA Glenn Research Center at Lewis Field about the possibility of applying Glenn's braided fiber preform seal as a thermal barrier to protect the O-ring seals. Glenn and Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and by using a braided carbon fiber thermal barrier that would resist any hot gases that the J-leg does not block.

An experimental investigation on the determination of ballistic properties and burning behavior of a composite solidpropellant for airbag application was conducted. The experimental results were obtained using a high-pressure optical strand burner. Steady-state burning rates were determined for a pressure range of 20.8-41.5 MPa and initial propellant temperatures of -30 to +80{sup o}C. For the pressure and temperature ranges tested, the temperature sensitivity was on the order of 1 x 10{sup -3} K{sup -1}. The pressure exponent was found to be a function of the initial propellant temperature and was 0.75 at 25{sup o}C. The activation energy and the pre-exponential factor of the Arrhenius equation are 2.735 kcal/mol and 15.06 cm/s, respectively. The pressure deflagration limit for this propellant was found to be in the range of 8.37-8.72 MPa. During combustion, small condensed-phase spherical particles were ejected from the burning surface. The size of the particles decreased with either increasing the pressure or the initial propellant temperature. For pressures below 41.4 MPa, average particle size was on the order of 900 {mu}m, and at 84.4 MPa, the bead size was much smaller, on the order of 300 {mu}m. A chemical analysis on these particles using both the ESEM and the X-ray diffraction method indicated that the material of the beads was mostly sodium chloride with a small amount of silicon-containing compounds. 6 refs., 17 figs.

Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

固体火箭发动机中,药柱的结构完整性直接关系到发动机的结构完整性和可靠性,而推进剂的力学性能对保持药柱结构完整性起着重要作用,也是决定推进剂寿命的重要指标.为了预估固体推进剂的力学性能,提高系统的可靠性,将遗传算法和神经网络相结合,建立了预估固体推进剂力学性能的遗传神经网络(GA-BP)模型.利用模型预测了某固体推进剂在不同温度、湿度和时间下的抗拉强度、延伸率、弹性模量变化情况,并与试验结果进行了比较.结果表明,模型预估精度高,泛化能力强,仿真计算与试验在结果上有很好的一致性.从而为固体火箭发动机的结构完整性研究提供可靠依据.%Solidpropellant grain structural integrity influences the structural integrity and reliability of solid rocket motor (SRM). Mechanical property of solidpropellant plays an important role in grain structural integrity, which is critical criterion of solidpropellant life. In order to predict mechanical property of solidpropellant, a new mechanical property prediction model for solid propollant was established by means of combination of 8enetic algorithm with neural network (GA-BP). Using above model, the mechanical proporty of a solidpropellant in conditions of different ternperature, humidity and time was predicted and compared with experiment results. The comparison results show high precision of the model and strong ability of generalization and with good consistency between prediction of model and experiment. The investigation provides reliable assistance for structural integrity research of SRM.

Foreign country has been successful in applying critical fluid extraction technology for demilitarization of solidpropellant and explosive. This paper introduced the process of critical fluid extraction technology through demilitarization of rocket motors containing ammonium percholrate (AP) composite propellants. The demilitarization process consists of four-step. Step one involves removing the AP propellant from the rocket motor. Step two extracts AP and separates the AP/liquid ammonia solution from binder residue. Step three recovers the AP by evaporating the ammonia. Step four condenses the ammonia vapor and recycles the liquid ammonia for a continuous removal/extraction operation.%通过对火箭发动机含高氯酸铵(AP)的复合推进剂用临界液氨销毁来介绍这一工艺方法。销毁工艺由四个连续的步骤构成：第一步采用临界液氨把含AP的复合推进剂从火箭发动机中取出；第二步萃取AP并将含AP的液氨和粘结剂残渣分离开；第三步从氨水中蒸发出AP；第四步将氨蒸气冷凝并重新循环使用。

In this paper,design and optimization technique of slotted tube grain for solid rocket motors has been discussed.In doing so,the design objectives and constraints have been set,geometric parameters identified,performance prediction parameters calculated,thereafter preliminary designs completed and finally optimal design reached.Geometric model for slotted tube grain configuration has been developed.Average thrust has been taken as the objective function with constraints of burning time,mass of propellant,fixed length and diameter of chamber case.Lumped parameter method has been used for calculating the performance prediction parameters.A set of preliminary designs has been completed and an analysis of these results conducted.Although all the preliminary results fulfill the design requirements in terms of objective function and constraints,however in order tO attain the optimal design,Sequen-tial quadratic programming optimization technique has been adopted.As the slotted tube grain ge-ometry is totally dependent upon various independent variables and each of these variables has a bearing on explicit characteristic of grain designing,hence affects of the independent variables on performance parameters have been examined,thus variation laws have been developed.Basing on the variation laws and the analysis of preliminary design results,upper and lower limits have been defined for the independent geometric variables and an initial guess provided for conducting optimi-zation.Resuhs attained exhibits that an optimal result has been attained and the value of objective function has been maximized.All the design constraint limits have also been met while ensuring sound values of volumetric loading fraction,web fraction and neutrality.This methodology of design and optimization of slotted tube grain for solid rocket motors can be used by engineers as a reference guide for actual design and engineering purposes.

An overview is presented of hybrid rocket propulsion systems whereby combining solids and liquids for launch vehicles could produce a safe, reliable, and low-cost product. The primary subsystems of a hybrid system consist of the oxidizer tank and feed system, an injector system, a solid fuel grain enclosed in a pressure vessel case, a mixing chamber, and a nozzle. The hybrid rocket has an inert grain, which reduces costs of development, transportation, manufacturing, and launch by avoiding many safety measures that must be taken when operating with solids. Other than their use in launch vehicles, hybrids are excellent for simulating the exhaust of solid rocket motors for material development.

Full Text Available In the framework of this research, the program Image Pro Plus was applied for determining the polymer–oxidizer interactions in HTPB-based composite propellants. In order to improve the interactions, different bonding agents were used, and their efficiency was analyzed. The determination of the quantity, area and radius of non-bonded oxidizer crystals is presented. The position of formed cracks in the specimen and their area has a great influence on the mechanical properties of composite propellant. The preparation of the composite propellant in order to enable the photographing of their structure by means of stereoscopic and metallographic microscopes with the digital camera is also described as well.

A flight investigation of two Nike-Deacon (DAN) two-stage solid-propellant rocket vehicles indicated satisfactory performance may be expected from the DAN meteorological sounding rocket. Peak altitudes of 356,000 and 350,000 feet, respectively, were recorded for the two flight tests when both vehicles were launched from sea level at an elevation angle of 75 degrees. Performance calculations based on flight-test results show that altitudes between 358,000 feet and 487,000 feet may be attained with payloads varying between 60 pounds and 10 pounds.

A Bayesian method is proposed which uses the safety margin to justify the design of high-reliability objects characterized by one-shot destructive functioning and by a high manufacturing cost. To validate the Bayesian reliability model, a program of overtests has been conducted with seven solidpropellant grains of six tons each. Each trial has been performed with a reduced safety margin leading to a design reliability of 0.5 at 60 percent confidence level. The reliability model gives reasonably good agreement with experimental results as the confidence level stabilizes, leading to the conclusion that it can be applied to estimate higher reliability levels (greater than 0.999).

It is known that the simulations of solid rocket motor internal flow field with AL-based propellants require complex multi-phase turbulent flow model. The objective of this study is to develop an advanced particulate multi-phase flow model which includes the effects of particle dynamics, chemical reaction and hot gas flow turbulence. The inclusion of particle agglomeration, particle/gas reaction and mass transfer, particle collision, coalescence and breakup mechanisms in modeling the particle dynamics will allow the proposed model to realistically simulate the flowfield inside a solid rocket motor. The Finite Difference Navier-Stokes numerical code FDNS is used to simulate the steady-state multi-phase particulate flow field for a 3-zone 2-D axisymmetric ASRM model and a 6-zone 3-D ASRM model at launch conditions. The 2-D model includes aft-end cavity and submerged nozzle. The 3-D model represents the whole ASRM geometry, including additional grain port area in the gas cavity and two inhibitors. FDNS is a pressure based finite difference Navier-Stokes flow solver with time-accurate adaptive second-order upwind schemes, standard and extended k-epsilon models with compressibility corrections, multi zone body-fitted formulations, and turbulence particle interaction model. Eulerian/Lagrangian multi-phase solution method is applied for multi-zone mesh. To simulate the chemical reaction, penalty function corrected efficient finite-rate chemistry integration method is used in FDNS. For the AL particle combustion rate, the Hermsen correlation is employed. To simulate the turbulent dispersion of particles, the Gaussian probability distribution with standard deviation equal to (2k/3)(exp 1/2) is used for the random turbulent velocity components. The computational results reveal that the flow field near the juncture of aft-end cavity and the submerged nozzle is very complex. The effects of the turbulent particles affect the flow field significantly and provide better

Full Text Available Ethylene-propylene diene terpolymer (EPDM-based insulation system is being globallyused for case-bonded solid rocket motors. A study was undertaken using EPDM as base polymer,blended with hypalon and liquid EPDM and filled with fibrous and non-fibrous fillers. Theseformulations were evaluated as rocket motor insulation system. The basic objective of the studywas to develop an insulation system based on EPDM for case-bonded applications. A series ofrocket motor insulator compositions based on EPDM, filled with particulate and fibrous fillerslike precipitated silica, fumed silica, aramid, and carbon fibres have been studied for mechanical,rheological, thermal, and interface properties. Compositions based on particulate fillers wereoptimised for the filler content. Comparatively, fumed silica was found to be superior as fillerin terms of mechanical and interface properties. Addition of fibrous filler (5 parts improved thepeel strength, and reduced the thermal conductivity and erosion rate. All the compositions wereevaluated for sulphur and peroxide curing. Superior mechanical properties were achieved forsulphur-cured products, whereas peroxide-cured products exhibited an excellent ageing resistance.Rocket motors were insulated with optimised composition and propellant cast, and the motorswere evaluated by conducting static test in end-burning mode.Defence Science Journal, 2008, 58(1, pp.94-102, DOI : http://dx.doi.org/10.14429/dsj.58.1628

using these equations derived by a formal averaging technique applied to the microscopic flow. These equations require a number of constitutive laws...disk (dimensions shown are from Chang and Howard [32]). acrylic, that allows cinematography of plasma flows and ignition events along the propellant

The vortex valve solid variable thrust motor is a new solidmotor which can achieve Vehicle system trajectory optimization and motor energy management. Numerical calculation was performed to investigate the influence of vortex chamber diameter, vortex chamber shape, and vortex chamber height of the vortex valve solid variable thrust motor on modulation performance. The test results verified that the calculation results are consistent with laboratory results with a maximum error of 9.5%. The research drew the following major conclusions: the optimal modulation performance was achieved in a cylindrical vortex chamber, increasing the vortex chamber diameter improved the modulation performance of the vortex valve solid variable thrust motor, optimal modulation performance could be achieved when the height of the vortex chamber is half of the vortex chamber outlet diameter, and the hot gas control flow could result in an enhancement of modulation performance. The results can provide the basis for establishing the design method of the vortex valve solid variable thrust motor.

The research effort outlined the application of a computer aided design (CAD)-centric technique to the design and optimization of solid rocket motor Finocyl (fin in cylinder) grain using simulated annealing.The proper method for constructing the grain configuration model, ballistic performance and optimizer integration for analysis was presented. Finoeyl is a complex grain configuration, requiring thirteen variables to define the geometry. The large number of variables not only complicates the geometrical construction but also optimization process. CAD representation encapsulates all of the geometric entities pertinent to the grain design in a parametric way, allowing manipulation of grain entity (web), performing regression and automating geometrical data calculations. Robustness to avoid local minima and efficient capacity to explore design space makes simulated annealing an attractive choice as optimizer. It is demonstrated with a constrained optimization of Finocyl grain geometry for homogeneous, isotropic propellant, uniform regression, and a quasi-steady, bulk mode internal ballistics model that maximizes average thrust for required deviations from neutrality.

Ballistic anomalies have frequently occurred during the firing of several solid rocket motors (SRMs) (Inertial Upper Stage, Space Shuttle Redesigned SRM (RSRM) and Titan IV SRM Upgrade (SRMU)), producing even relevant and unexpected variations of the SRM pressure trace from its nominal profile. This paper has the purpose to provide a numerical analysis of the following possible causes of ballistic anomalies in SRMs: an inert object discharge, a slag ejection, and an unexpected increase in the propellant burning rate or in the combustion surface. The SRM configuration under investigation is an aft-finocyl SRM with a first-stage/small booster design. The numerical simulations are performed with a quasi-one-dimensional (Q1D) unsteady model of the SRM internal ballistics, properly tailored to model each possible cause of the ballistic anomalies. The results have shown that a classification based on the head-end pressure (HEP) signature, relating each other the HEP shape and the ballistic anomaly cause, can be made. For each cause of ballistic anomalies, a deepened discussion of the parameters driving the HEP signatures is provided, as well as qualitative and quantitative assessments of the resultant pressure signals.

为获得羽烟对激光透过率的影响,用烟箱法对2种配方的缩比发动机羽烟在1.06 μm、10.6 μm激光波段的透过率进行测试.采用1.064 μm激光调制发射、接收、数据采集系统对1.06 μm激光波段烟雾透过率测试;用黑体、光谱辐射计、数据采集系统可测出2 μm ～13 μm 的光学透过率,从中提出10.6 μm激光波段烟雾透过率,得到不同推进剂配方、不同烟雾浓度情况下10.6 μm光波和1.06 μm光波的烟雾透过率测试数据.烟箱1.8 m烟道上的测试数据表明:配方2推进剂优于配方1推进剂,10.6 μm光波的烟雾透过率96%～97%大于1.06 μm光波的烟雾透过率92%～93%.%The main performance of hyper-velocity kinetic missile and its solidpropellants is described. The mechanism of laser attenuation caused by the plume of solid rocket motor is analyzed. The plume transmittance of two solid-propellant formulations was tested by the aid of a smoke-box when the laser wavebands were 1.06 μm and 10.6 μm. The plume transmittance in 1.06 μm laser was tested with 1.064 μm laser modulation emitting, receiving and data acqui-sition system. The plume transmittance of 10.6 μm laser was derived after the optical transmittance of 2 μm ～13 μm laser was tested with the blackbody radiation source, spectroradiometer and data acquisition system. The tested data indicates that formulation 2 is better than formulation 1, and the plume transmittance 96%～97% of 10.6 μm laser is higher than that 92%～93% of 1.06 μm laser.

针对复合固体推进剂的老化问题，探讨了复合固体推进剂老化的主要因素；从复合固体推进剂组分的影响、环境湿度、储存温度等方面分析了其影响复合固体推进剂老化的机理；综述了复合固体推进剂贮存寿命的几种预估方法：力学性能法、阿累尼乌斯方程式法、凝胶含量法、傅里叶红外光谱法和动态粘弹法，并对这些研究方法的内容和结果可信度进行了分析。最后对复合固体推进剂老化研究的发展趋势进行了展望。%In the view of the composite solidpropellant ageing,the main factors of the composite solidpropellant aging were discussed and aging mechanism of the composite solidpropellant was analyzed from the influence of the component of solidpropellant,environment humidity and storing temperature,several forecast methods of composite solidpropellant storage life was overviewed:the methods of mechanical properties, activation energy, gel content, Fourier infrared spectrum analysis and dynamic mechanical analysis. And simultaneously the main content of these methods and the reliability of these results were analyzed. Finally,aging research trends of the composite solidpro-pellant were discussed.

Downsizing has become essential in the development of motors recently in order to meet requirements related to environmental issues such as resource depletion and energy shortages. To fulfill these demands, we propose downsizing techniques for a conventional motor where a solid insulation system is hired and a large amount of insulation resin was employed to replace structural metal materials used in conventional motors. This paper deals with a feasibility study of a solid insulation system applied downsized motor. The optimal heat design, mechanically stable structure, and proper insulation resin are proposed and discussed. A prototype motor was manufactured, and results of a basic insulation test, temperature rise test and noise test are presented.

Full Text Available Poisson's ratio of hydroxy-terminated polybutadine (HTPB-based composite propellant is estimated from uni-axial tensile testing. Double dumbbell specimens as per ASTM D638 type IV standard were used and Poisson's ratio at break, obtained by change in volume of specimen, was calculated as approximately 0.25. It was also observed that Poisson's ratio is different along different lateral directions of the propellant specimen. Poisson's ratios in two orthogonal directions perpendicular to longitudinal axis were calculated as 0.17 and 0.30. As ASTM specimen has rectangular cross-section of approximate size 6 mm x 4 mm, the directional behaviour of Poisson's ratio may be attributed to initial dimensions. Prismatic propellant specimen with square cross-section and of 115 mm x 6 mm x 6 mm dimension do not show any variation wrt Young's modulus,tensile strength, and percentage elongation as compared to ASTM specimen. Directional behaviour of Poisson's ratio with almost similar numerical value was again observed, thus ruling out dependence of this behaviour on different initial dimensions of propellant cross-section. Further, Poisson's ratio varies linearly with strain even in linear portion of stress-strain curve in uni-axial tensile testing. The rate of reduction of Poisson's ratio with increase in strain is slower in linear region and it accelerates after dewetting due to formation of vacuoles. Variation of Poisson's ratio with strain has two different slopes in linear (slope = 0.3165 and nonlinear regions (slope = 0.61364. Numerical value of slope for variation of Poisson's ratio with strain almost doubles after dewetting. It must be noted that no change in volume does not necessarily indicate constant Poisson's ratioequal to 0.5. Composite propellants behave as compressible material in most of the regions and near-failure region or at higher strains; Poisson's ratio is not anywhere near to 0.5, instead it is near 0.25.Defence Science

The program planning acquisition functions for the development of the solidpropellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.

This work describes the evaluation of various solid phase micro-extraction (SPME) fibre types for the detection of compounds originating from particles of unburned propellant powders. These compounds may also be found in association with organic gunshot residues (OGSR). Seven SPME fibres were assessed based on their ability to extract the compounds of interest (diphenylamine (DPA), 4-nitrodiphenylamine (4-NDPA), ethyl centralite (EC), nitroglycerin (NG) and dibutyl phthalate (DBP)) from four ammunition types across three calibres (9 mm, 5.56 mm and 7.62 mm). Extracts were analysed by gas chromatography/mass spectrometry (GC/MS). Results indicated that the 65 μm polydimethylsiloxane/divinylbenzene (PDMS/DVB) was the most suitable fibre type for the extraction of these compounds across the ammunition types tested. Optimal extraction time parameters were also assessed with a 35-min period determined to be suitable. A number of previously unreported considerations for extracting propellant powders and potentially OGSR related materials are discussed.

WP 4gw«’WUK_u.,HU!«J»B^!PB!*-pi!^p5ggB5,^.IJll.l II Hill JMflH» !,S--*"JU’j*llfgg V. D. Organometallic Compounds E. Enthalpy and Gas Composition...zone near the burning surface of the propellent where the reactive species can be neutralised more effectively. B. Char Formers Several phisphorus... Enthalpy and Gas Composition Measurements of Coolants and Binders* Two experimental techniques were used to measure the enthalpy changes that are

Full Text Available In the light of the recent Chinese aggression, the modernization of our defence forces has become a vital necessity. In this paper, one aspect of this problem has been considered in detail viz. the development of an indigenous rocket and missile force. While it is true that many factors e.g. rocket motors, propellants, guidance systems etc. are involved, it is also true that a start in one area will act as an impetus to developments in the other fields. Solid/liquid propellant and oxidizer systems have been considered, the properties of solid and liquid propellants evaluated and on the basis of such comparison, it has been concluded that effort concentrated on the development of liquid propellants will be well expended. Liquid propellant/oxidizer systems have been compared amongst themselves and it has been concluded that the hydrazine fuels oxidized by RFNA/WFNA/IRFNA would represent systems fulfilling the country's immediate military needs best. The availability of raw materials for the manufacture of hydrazine fuels (and also of some solidpropellants has been considered and it is shown that the necessary raw materials are available in sufficient quantities to support an indigenous propellants industry.

Full Text Available With constant improvements, the conventional solidpropellants for guns have almost reached their limit in performance. Liquid gun propellants are promising new comers capable of surpassing these performance limits and have numerous advantages over solidpropellants. A method has been worked out to predict the internal ballistics of a liquid propellant gun and illustrated in a typical application.

A physicomathematical model of combustion of a metallized composite solidpropellant based on ammonium perchlorate has been presented. The model takes account of the thermal effect of decomposition of a condensed phase (c phase), convection, diffusion, the exothermal chemical reaction in a gas phase, the heating and combustion of aluminum particles in the gas flow, and the velocity lag of the particles behind the gas. The influence of the granulometric composition of aluminum particles escaping from the combustion surface on the linear rate of combustion has been investigated. It has been shown that information not only on the kinetics of chemical reactions in the gas phase, but also on the granulometric composition of aluminum particles escaping from the surface of the c phase into the gas, is of importance for determination of the linear rate of combustion.

引入了光致正电子湮灭分析方法 （PIPA,Photon Induced Positron Analysis）,对HTPB固体推进剂老化性能进行检测研究。介绍了PIPA的原理、PIPA试验平台的搭建以及PIPA的数值处理方法,并用511KeV能谱的FWHM值ε（半高宽,Full Width Half Maximum）表征HTPB固体推进剂老化引起的微观变化,所得结论与固体推进剂的常规测试一致,证明了PIPA用于固体推进剂老化性能无损检测的可行性。%To study the aging character of HTPB solidpropellant,PIPA（Photon Induced Positron Analysis） was presented.The positrons created by the PIPA process are formed throughout the solidpropellant.The principle of PIPA was introduced along with the built-up of the PIPA system and the arithmetic method.The information of HTPB solidpropellant micro change due to aging was extracted from the FWHM of 511keV spectra.There is positive correlation between the results of experiments and regular tests,indicating that PIPA is a feasible nondestructive detection method for solidpropellant aging evaluation.

This paper presents the production scheme for heat- and erosion-protective carbon plastic materials for heat shield elements of solid-propellant nozzles. Attention is also given the method of manufacturing adhesive joint assemblies, and the production scheme is included.

The Five-Segment Engineering Test Motor (ETM-3) is an extended length reusable solid rocket motor (RSRM) intended to increase motor performance and internal environments above the current four-segment RSRM flight motor. The principal purpose of ETM-3 is to provide a test article for RSRM component margin testing. As the RSRM and Space Shuttle in general continue to age, replacing obsolete materials becomes an ever-increasing issue. Having a five-segment motor that provides environments in excess of normal opera- tion allows a mechanism to subject replacement materials to a more severe environment than experienced in flight. Additionally, ETM-3 offers a second design data point from which to develop and/or validate analytical models that currently have some level of empiricism associated with them. These enhanced models have the potential to further the understanding of RSRM motor performance and solid rocket motor (SRM) propulsion in general. Furthermore, these data could be leveraged to support a five-segment booster (FSB) development program should the Space Shuttle program choose to pursue this option for abort mode enhancements during the ascent phase. A tertiary goal of ETM-3 is to challenge both the ATK Thiokol Propulsion and NASA MSFC technical personnel through the design and analysis of a large solid rocket motor without the benefit of a well-established performance database such as the RSRM. The end result of this undertaking will be a more competent and experienced workforce for both organizations. Of particular interest are the motor design characteristics and the systems engineering approach used to conduct a complex yet successful large motor static test. These aspects of ETM-3 and more will be summarized.

Full Text Available A general numerical model based on the Zeldovich-Novozhilov solid-phase energy conservation result for unsteady solid-propellant burning is presented in this paper. Unlike past models, the integrated temperature distribution in the solid phase is utilized directly for estimating instantaneous burning rate (rather than the thermal gradient at the burning surface. The burning model is general in the sense that the model may be incorporated for various propellant burning-rate mechanisms. Given the availability of pressure-related experimental data in the open literature, varying static pressure is the principal mechanism of interest in this study. The example predicted results presented in this paper are to a substantial extent consistent with the corresponding experimental firing response data.

Full Text Available Calorimetric value and ignition temperature of carboxy terminated polybutadiene/ammonium perchlorate propellant decrease during accelerated ageing. The behaviour has been explained on account of binder loss and condensed phase reactions in the propellant matrix.

One of the most critical areas for inspection in the Space Shuttle Solid Rocket Motors is the bond between the steel case and rubber insulation in the region of the field joints. The tang-and-clevis geometry of the field joints is sufficiently complex to prohibit the use of resonance-based techniques. One approach we are investigating is to interrogate the steel-insulation bondline in the tang and clevis regions using surface-travelling waves. A low-frequency contact surface wave transmitting array transducer is under development at our laboratory for this purpose. The array is placed in acoustic contact with the steel and surface waves are launched on the inside surface or the clevis leg which propagate along the steel-insulation interface. As these surface waves propagate along the bonded surface, the magnitude of the ultrasonic energy leaking into the steel is monitored on the outer surface of the case. Our working hypothesis is that the magnitude of energy received at the outer surface of the case is dependent upon the integrity of the case-insulation bond, with less attenuation for propagation along a disbond due to imperfect acoustic coupling between the steel and rubber. Measurements on test specimens indicate a linear relationship between received signal amplitude and the length of good bend between the transmitter and receiver, suggesting the validity of this working hypothesis.

Asbestos free solidmotor internal insulation samples were tested at the MSFC Hyperthermal Facility. Objectives of the test were to gather data for analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Instrumentation included use of total calorimeters, thermocouples, and a surface pyrometer for surface temperature measurement. Post-test sample forensics involved measurement of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero-thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

Solid rocket motors (SRMs) typically use nozzle materials which are required to maintain their shape as well as insulate the underlying support structure during the motor operation. In addition, SRMs need internal insulation materials to protect the motor case from the harsh environment resulting from the combustion of solidpropellant. In the nozzle, typical materials consist of high density graphite, carbon-carbon composites and carbon phenolic composites. Internal insulation of the motor cases is typically a composite material with carbon, asbestos, Kevlar, or silica fibers in an ablative matrix such as EPDM or NBR. For both nozzle and internal insulation materials, the charring process occurs when the hot combustion products heat the material intensely. The pyrolysis of the matrix material takes away a portion of the thermal energy near the wall surface and leaves behind a char layer. The fiber reinforcement retains the porous char layer which provides continued thermal protection from the hot combustion products. It is of great interest to characterize both the total erosion rates of the material and the char layer thickness. By better understanding of the erosion process for a particular ablative material in a specific flow environment, the required insulation material thickness can be properly selected. The recession rates of internal insulation and nozzle materials of SRMs are typically determined by testing in some sort of simulated environment; either arc-jet testing, flame torch testing, or subscale SRMs of different size. Material recession rates are deduced by comparison of pre- and post-test measurements and then averaging over the duration of the test. However, these averaging techniques cannot be used to determine the instantaneous recession rates of the material. Knowledge of the variation in recession rates in response to the instantaneous flow conditions during the motor operation is of great importance. For example, in many SRM configurations

A series of tests involving the combustion of solid and liquid propellants used to fuel the Titan 2, Titan 4, and Delta 2 launch vehicles was performed. The purpose of these tests was to evaluate the nature and amounts of combustion gases from reacting these propellants in various proportions, and to apply the derived data to predicting toxic chemical emissions arising from a launch vehicle explosion. Propellants tested in this study included Aerozine-50 and nitrogen tetroxide (liquid propellants used in the Titan 2 and Titan 4 launch vehicles), PBAN solidpropellant (used on the Titan 4 solid rocket motor), RP-1 and liquid oxygen (liquid propellants used to fuel the Delta 2 launch vehicle), and the Castor IVA solid rocket propellant used on the Delta 2 first stage engine. Tests were conducted in a 150-liter stainless steel combustion chamber in air at nominal pressure (0.8 atmospheres at Denver barometric conditions). Measurements of the chamber gas temperature and internal pressure were taken and gas samples were withdrawn and analyzed for expected combustion gases, unreacted propellants, organic vapors, and oxygen reacted from the air. A stainless steel witness plate was used to collect condensates which formed during the course of the propellant combustion tests. Results of this study suggest significantly different chemical fates for some of the rocket propellants than those predicted by chemical theory only. A description of the test parameters, results, and application to source predictions is presented.

The emerging field of soft robotics relies on soft, stimuli-responsive materials to enable load transport, manipulation, and mobility in complex unconstrained environments. These materials often need to replicate biological functionality such as muscle contractions and flexibility. Here we demonstrate a soft actuator prototype based on thermosensitive PNIPAAM hydrogels that can transport and manipulate objects. A hollow cylindrical hydrogel was selectively heated and cooled with Peltier devices to yield a traveling wave of shrinking and swelling akin to intestinal peristalsis. A 4 mm diameter bead was placed inside the cylinder and propelled 19.5 mm, equal to distance traveled by the peristaltic wave. We derived conditions that enable peristaltic transport as a function of transporter-cargo design parameters. We conclude that hydrogel-based peristaltic manipulators covering 2 orders of magnitude in stiffness (1-10(2) kPa) could transport cargo spanning 4 orders of magnitude in size (μm-m).

The combustion of NHClO4 composite propellants was studied between 15 and 3000 psi. The emphasis was on determining the mechanisms by which the fuel components influence the burning rate of the composites. All combustion experiments were performed with pressed powder strands. The fuels affected the combustion mechanism of the composite through their effect on the oxidizer decomposition mechanism, the composite surface temperature, and by reacting heterogeneously with the oxidizer on the surface of the composite. The results indicate that the fuel component has a significant effect on the surface temperature in both catalyzed and uncatalyzed composites. The data suggest that the contribution, to the overall combustion reaction, of heterogeneous reactions increases with increasing pressure. There is also evidence that even in steady state combustion the ignition of individual oxidizer particles is an important parameter in determining the combustion characteristics of the composite.

The results of an investigation are presented on the applicability of a microwave Doppler shift technique for directly determining propellant response functions over the desired frequency range. The investigation consisted of three phases. In Phase 1 the validity of the technique was established by comparing measured pressure-coupled response function data to existing data from T-burners and rotating valve tests. In Phase 2 a new microwave burner-pressure modulation system capable of achieving frequencies and mean chamber pressures of at least 1500 Hz and 10.5 MPa (1500 psia), respectively, was developed. During Phase 3 test firings are being carried out to define the frequency limit, response function resolution, and precision of the new design.

cleaners, such as a terpene /surfactant mixture. Organic solvents generally are effective in removing heavy oils and greases; however, they may pose...occur. Small decreases in DTA exotherms were observed indicating that some solvent was absorbed into the propellant; however, no significant safety

The research status of the crack extension in solidpropellant at home and abroad are introduced. Experimental study and theoretical analysis of the crack extension in solidpropellant are summarized. The advantages and disadvantages of such study are analyzed. The influences of various factors on crack extension in solidpropellant are generalized. The results show that the tension ma- chine and high-speed photography are main means for studying the mechanics characteristic of pro- pellant with crack, and the combustion property of propellant and the extension of crack; the working pressure in combustion chamber, pressure boost rate, style and size of crack, and burning rate of pro- pellant arc the important factors for the crack extension in solidpropellant. It is pointed out that it is necessary to research the quantitative relationship between various factors and crack extension.%介绍国内外关于固体推进剂裂纹扩展的研究现状，总结固体推进剂裂纹扩展的试验研究和理论分析方法，归纳影响固体推进剂裂纹扩展的各种因素。分析认为：采用拉伸装置研究有裂纹推进剂力学性能和利用高速摄影系统研究裂纹燃烧和扩展情况是当前主要的研究手段；发动机燃烧室内压力、升压速率、裂纹形状尺寸和推进剂燃速是装药裂纹扩展的重要因素；必须进一步开展各因素和裂纹扩展的定量关系研究。

The thermal stability of perchlorate composite propellants was studied at 135 and 170 C. The experimental efforts were concentrated on determining the importance of heterogeneous oxidizer-fuel reactions in the thermal degradation process. The experimental approach used to elucidate the mechanisms by which the oxidizer fuel composites thermally degrade was divided into two parts: (1) keeping the fuel constant and varying the nature of the oxidizers, and (2) holding the oxidizer constant and varying the fuel components. The fuel component primarily utilized in the first phase was polyethylene. Oxidizers included KClO4, KClO3, NH4ClO4 and NH4ClO4 doped with materials such as chlorate, phosphate and arsenate. In the second phase the oxidizer used was primarily NH4ClO4 while the fuels included saturated and unsaturated polybutadiene prepolymers and a series of bonding agents. Techniques employed in the current study include thermogravimetric measurements, differential thermal analysis, infrared, mass spectrometry, electron microscopy, and appropriate wet chemical analysis.

National Aeronautics and Space Administration — Solid rocket motor cases are subject to a variety of external environmental and loading conditions from cradle-to-grave. These conditions can significantly impact...

National Aeronautics and Space Administration — Small satellites such as CubeSats are in need of responsive propulsion, but are limited due to their size. Though single pulse, AP/HTPB fueled solid rocket motors...

Thrust-terminating system conceived to reduce thrust of solid-propellant rocket motor in controlled manner such that thrust loads not increased or decreased beyond predictable levels. Concept involves explosively cutting opposing venting pairs in case of rocket motor above nozzles to initiate venting of chamber and reduction of thrust. Vents sized and numbered to control amount and rate of reduction in thrust.

介绍了固体推进剂用粘合剂的研究进展,着重介绍了丁羟类、叠氮类、HTPE、NEPE以及硝基和硝酸酯类粘合剂的合成、工艺以及性能研究情况,并对未来固体推进剂用粘合剂的发展进行了展望.%Research results of the binders for in solidpropellants mainly including the synthesis and performance of hydroxyl terminated polubutadiene (HTPB) binder.azido based energetic binders,hydroxyl terminated polyrther (HTPE) binder.nitro and nitrato group containing binders were introduced.And the prospective development of binders for solidpropellants was discussed.

A comprehensive numerical analysis has been carried out to study the unsteady flowfields in a simulated rocket-motor environment. The model is based on the time-dependent compressible Navier-Stokes equations with a two-equation turbulence closure scheme. Various important aspects of the coupling between acoustic oscillations and mean flowfields, including flow reversal, modification of transport properties, etc., are addressed. Results indicate that multi-dimensional effects play important roles in determining local flow structures and wave characteristics. In much of the domain, acoustic velocity nodal points are observed in the near-wall region. The classical one-dimensional theory fails to describe several important mechanisms associated with velocity-induced flow instabilities.

This paper presents the development stage of a communication module for a solidpropellant mid-power rocket model. The communication module was named. Simple-1 and this work considers its design, construction and testing. A rocket model Estes Ventris Series Pro II® was modified to introduce, on the top of the payload, several sensors in a CanSat form factor. The Printed Circuit Board (PCB) was designed and fabricated from Commercial Off The Shelf (COTS) components and assembled in a cylindrical rack structure similar to this small format satellite concept. The sensors data was processed using one Arduino Mini and transmitted using a radio module to a Software Defined Radio (SDR) HackRF based platform on the ground station. The Simple-1 was tested using a drone in successive releases, reaching altitudes from 200 to 300 meters. Different kind of data, in terms of altitude, position, atmospheric pressure and vehicle temperature were successfully measured, making possible the progress to a next stage of launching and analysis.

The effect of increasing solid phase concentration on the morphology and flocculation rate of model aerosol suspensions has been investigated. Suspensions of micronized salbutamol sulphate and lactose in trichlorotrifluoroethane (P113) were studied under conditions of increasing shear stress. By use of image analysis techniques, measurement of aggregate size, fractal dimension and rate of aggregation was performed. The effect of the surfactant sorbitan monooleate on morphology and flocculation rate was also studied. Increased solid phase concentration caused an increase in the rate of aggregation and average aggregate size at a given value of shear stress. Surfactant addition retarded the aggregation rate, and caused a shift from a diffusion-limited cluster aggregation to a reaction-limited cluster aggregation mechanism. The aggregate profiles showed a corresponding change from rugged and crenellated without surfactant, to increasingly smooth and Euclidian with increasing surfactant concentration. The morphological changes were characterized by a decrease in the average boundary fractal dimension which also correlated well with the corresponding reduction in aggregation rate.

This patent describes an azido-based solidpropellant composition having an improved burning rate comprising: a high energy plasticizer of tris-1,2,3(bis(1,2-difluoroamino)ethoxy)propane in an amount from about 24 to about 30 weight percent of the propellant composition; a curative and crosslinking agent of 4,5-epoxycyclohexylmethyl 4'5'-epoxycyclohexylcarboxylate in an amount from about 0.75 to about 1.5 weight percent of the propellant composition; a carboranyl burning rate catalyst of carboranyl-methyl propionate in an amount from about 2 to about 6 weight percent of the propellant composition; graphite linters of about 100 micrometers lengths in an amount from about 1 to about 3 weight percent of the propellant composition; aluminum powder in an amount from about 10 to about 12 weight percent of the propellant composition; aluminum flake in an amount from about 0.5 to about 2 weight percent of the propellant composition; ammonium perchlorate of about 0.9 micrometer diameter in an amount from about 46 to about 52 weight percent of the composition; a processing aid of lecithin in an amount from about 0.1 to about 0.2 weight percent of the propellant composition; and a binder of 2-azidoethyl acrylateacrylic acid copolymer in an amount from about 3 to about 8 weight percent of the propellant composition.

Solidpropellant consists of incendiary and oxidant. Solidpropellant can combust intensely without oxygen. Aiming at the main causes of propellant inflammation during the rolling manufacture, the effect of temperature and moisture content on hot-spot formation of solidpropellant was investigated. Further more, the technical safe measures for solidpropellant manufacture. Aiming at the extinguishing problem that the fire fighting equipment being used in the solidpropellant industry can not effectively put out the solidpropellant fire, the ideas and methods of the solidpropellant fire extinguishing were basically brought forward by the analysis of the solidpropellant fire mechanism. Experimented results indicated that the choose of spray pressure, spray bore diameter, distance from spray to fire and fire time in anticipation plays an important role in extinguishing the solidpropellant fire. As a result, the ideas and methods of the high-speed water jet technology for extinguishing the solidpropellant fire.%固体推进剂的组分中既有氧化剂又有燃烧剂,在没有外界氧气的情况下,也能发生燃烧爆炸事故.针对固体推进剂在压延生产中产生燃烧的主要原因,研究了水分和温度对固体推进剂热点形成的影响,提出了防火安全技术措施.针对部分火炸药企业的消防设施不能有效扑灭固体推进剂火灾和阻止燃爆事故的情况,通过对固体推进剂燃烧灭火机理的分析,采用选择不同的喷水压力、喷头孔径、喷头与火焰区的距离、预先燃烧时间等因素,对固体推进剂进行燃烧灭火效果的实验研究,提出了固体推进剂火灾的高速喷水灭火技术的思路和措施.

Laser propulsion has been developed as a suitable small thruster technology for the attitude control of micro and nano class satellites. Laser-based thrusters meet the satellite design criteria for being of light weight and cost effective, because they do not require fuel storing and oxidizer for combustion. Also, thrust control by laser propulsion is achieved fairly easy. In this paper, we consider the confinement of plasma expansion by a gel-type liquid material, which results in the enhancement of the thrust for propulsion. We also present our attempts to resolve some known issues regarding laser ablation of solid and liquid targets. The level of thrust is quantified via the momentum coupling coefficient, which was experimentally measured using a ballistic pendulum system. We have discovered that the laser ablation confinement by the gel-type medium results in 2.3 times more enhanced driving force as compared to the water confinement. A proof of performance is demonstrated for using gel-type material for generating propulsion, and material characterization for optimal thrust performance is presented.

Laser propulsion has been developed as a suitable small thruster technology for the attitude control of micro and nano class satellites. Laser-based thrusters meet the satellite design criteria for being of light weight and cost effective, because they do not require fuel storing and oxidizer for combustion. Also, thrust control by laser propulsion is achieved fairly easy. In this paper, we consider the confinement of plasma expansion by a gel-type liquid material, which results in the enhancement of the thrust for propulsion. We also present our attempts to resolve some known issues regarding laser ablation of solid and liquid targets. The level of thrust is quantified via the momentum coupling coefficient, which was experimentally measured using a ballistic pendulum system. We have discovered that the laser ablation confinement by the gel-type medium results in 2.3 times more enhanced driving force as compared to the water confinement. A proof of performance is demonstrated for using gel-type material for generating propulsion, and material characterization for optimal thrust performance is presented. (orig.)

Analytical Solution Utilizing the geometry and dimensions provided in Figure 2, and considering that the motor is axisymmetric, the burn...distribution is unlimited. PA# Analytical Solution Motor Geometry Burn Surface Area, Ab 118,996 mm2 Throat Area, At 531 mm2 • From previous equations...FEA) is explored. A propellant flap in a cross flow is analyzed. Comparisons are made between an analytical solution, a solely CFD solution, a manual

SEP is in charge since 1989 of the ERINT-1 motor case and nozzle. The stringent missile weight and volume requirements coupled with the specification to provide an aerodynamically stable configuration over a very large Mach number range led to the need to develop a high-performance composite motor case. Development of this SRM case presented a variety of technical challenges that were solved by an original design: (1) integral skirts, high bending stiffness, and bending loads are required; (2) high temperature composite stiffness and loads are required up to 160 C; (3) integral fin lugs attachments high aerodynamic loading is required on fin lugs; (4) enclosed fore dome; and (5) aft-pinned joint: a large rear opening is required to cast the propellant. Structural testing in ultimate conditions confirmed the soundness of the design. Positive safety margins were demonstrated on both internal pressure and mechanical loads requirements.

butadiene Acrylic Acid Propellants 14 10 807 AP + 20% PBAA Data Plotted as (p/r) vs (p2/3) 14 11 DB and CMDB Propellant Data Plotted as (p/r) vs (p2/3...1.2% stabilizer causes a striking differ- ) ence in its burning behavior. This composite-modified double-base ( CMDB ) pro- pellant burns like a normal...dominated by a granular diffusion flame mechanism. It is to be noted that since the binder in a CMDB propellant is itself a monopropellant, there is no

Full Text Available Dual-thrust mode is adopted in solidpropellant rocket propulsion through tailoring of burning area, nozzle, rocket motor chamber, propellant type, multiple propellant blocks. In the present study, mathematical formulation has been evolved for generation of burning surface area with web burnt for a simple central blind hole in a solid cylindrical propellant geometry with proper partial inhibition on external and lateral surfaces. The burn-back equation has been validated by static firing and parametric study was conducted to understand effect of various control geometrical parameters. The system is utilised for high volumetric loading, single propellant, single composition, single-chamber, single nozzle dual-thrust mode of burning profiles in rocket application.Defence Science Journal, 2011, 61(2, pp.165-170, DOI:http://dx.doi.org/10.14429/dsj.61.41

The Reusable Solid Rocket Motor (RSRM) represents the largest solid rocket motor (SRM) ever flown and the only human-rated solidmotor. High reliability of the RSRM has been the result of challenges addressed and lessons learned. Advancements have resulted by applying attention to process control, testing, and postflight through timely and thorough communication in dealing with all issues. A structured and disciplined approach was taken to identify and disposition all concerns. Careful consideration and application of alternate opinions was embraced. Focus was placed on process control, ground test programs, and postflight assessment. Process control is mandatory for an SRM, because an acceptance test of the delivered product is not feasible. The RSRM maintained both full-scale and subscale test articles, which enabled continuous improvement of design and evaluation of process control and material behavior. Additionally RSRM reliability was achieved through attention to detail in post flight assessment to observe any shift in performance. The postflight analysis and inspections provided invaluable reliability data as it enables observation of actual flight performance, most of which would not be available if the motors were not recovered. RSRM reusability offered unique opportunities to learn about the hardware. NASA is moving forward with the Space Launch System that incorporates propulsion systems that takes advantage of the heritage Shuttle and Ares solidmotor programs. These unique challenges, features of the RSRM, materials and manufacturing issues, and design improvements will be discussed in the paper.

The aerodynamic effects of plumes from hot combustion gases in the presence of a transonic external flow field were measured to advance plumes simulation technology, extend a previously acquired data base, and provide data to compare with the effects observed using cold gas plumes. A variety of underexpanded plumes issuing from the base of a strut-mounted ogive-cylinder body were produced by combusting solidpropellant gas generators. The gas generator fired in a short-duration mode (200 to 300 msec). Propellants containing 16 percent and 2 percent A1 were used, with chamber pressures from 400 to 1800 psia. Conical nozzles of 15 deg half-angle were tested with area ratios of 4 and 8. Pressures were measured in the gas generator combustion chamber, along the nozzle wall, on the base, and along the body rear exterior. Schlieren photographs were taken for all tests. Test data are presented along with a description of the test setup and procedures.

It is reported that the neccessary technology has been developed and demonstrated for the manufacture of heat-sterilizable solidpropellants which meet specific ballistic goals. It is shown that: (1) phosphate doping of ammonium perchlorate significantly enhances the thermal stability of the substance; (2) grinding the ammonium perchlorate to reduce particle size further increases thermal stability; and (3) unsaturated polymers such as the polybutadienes can be successfully used in a heat-sterilizable propellant system. Among the topics considered by the study are oxidizers, dopants, binders, and the thermal cycling of 70 lb and 600 lb propellant grains.

Rocket propulsion is an important travel method for space exploration and national defense, rockets needs to be able to withstand wide range of operation environment and also stable and precise enough to carry sophisticated payload into orbit, those engineering requirement makes rocket becomes one of the state of the art industry. The rocket family have been classified into two major group of liquid and solid rocket based on the fuel phase of liquid or solid state. The solid rocket has the advantages of simple working mechanism, less maintenance and preparing procedure and higher storage safety, those characters of solid rocket make it becomes popular in aerospace industry. Aluminum based propellant is widely used in solid rocket motor (SRM) industry due to its avalibility, combusion performance and economical fuel option, however after aluminum react with oxidant of amonimum perchrate (AP), it will generate liquid phase alumina (Al2O3) as product in high temperature (2,700˜3,000 K) combustion chamber enviornment. The liquid phase alumina particles aggromorate inside combustion chamber into larger particle which becomes major erosion calprit on inner nozzle wall while alumina aggromorates impinge on the nozzle wall surface. The erosion mechanism result nozzle throat material removal, increase the performance optimized throat diameter and reduce nozzle exit to throat area ratio which leads to the reduction of exhaust gas velocity, Mach number and lower the propulsion thrust force. The approach to avoid particle erosion phenomenon taking place in SRM's nozzle is to reduce the alumina particle size inside combustion chamber which could be done by further breakup of the alumina droplet size in SRM's combustion chamber. The study of liquid breakup mechanism is an important means to smaller combustion chamber alumina droplet size and mitigate the erosion tack place on rocket nozzle region. In this study, a straight two phase air-water flow channel experiment is set up

We report the synthesis of altitudinal molecular motors that contain functional groups in their rotor part. In an approach to achieve dynamic control over the properties of solid surfaces, a hydrophobic perfluorobutyl chain and a relatively hydrophilic cyano group were introduced to the rotor part of the motors. Molecular motors were attached to quartz surfaces by using interfacial 1,3-dipolar cycloadditions. To test the effect of the functional groups on the rotary motion, photochemical and thermal isomerization studies of the motors were performed both in solution and when attached to the surface. We found that the substituents have no significant effect on the thermal and photochemical processes, and the functionalized motors preserved their rotary function both in solution and on a quartz surface. Preliminary results on the influence of the functional groups on surface wettability are also described.

Full Text Available The article gives a detailed molecular-kinetic theory of the Mobius propeller functioning and shows the implementation of its computer modelling in different exploitation conditions. The mechanisation of the Mobius propeller can be carried out in such a way that, under certain conditions, it enables using this type of propellers as a heat pump. The developed theory of the Mobius propeller functioning has been experimentally verified in laboratory conditions.

To better simulate the process that mesoscopic damage evolves into macro cracks of solidpropellants and the effect of this process on the nonlinear mechanical properties of composite solidpropellants, the molecular dynamics method was adopted to create particle packing models of solidpropellants. Based on the surface-based cohesive approach, interfacial damages between par-ticles and the bind are created. The finite element method was adopted to compute the packing models with damage and the effect of solid content and interfacial damage on mesoscopic damage morphology and mechanical properties of solidpropellant has been re-searched by comparing the numerical simulation results. The results show that interfacial damage always gathers among a few parti-cles when the solid content is low. With the increasing of solid content, particles involved in the process of interfacial damage aggre-gation increase more and more. The nonlinear mechanical properties of composite solidpropellants, which are affected greatly by in-terfacial damage between particles and the binder, can not be ignored.%为准确模拟固含量不同时复合固体推进剂细观损伤产生、演化、聚合至宏观裂纹形成的过程，及该过程对复合固体推进剂非线性力学性能的影响，采用分子动力学方法建立了复合固体推进剂颗粒夹杂模型，根据Surface-based cohesive方法，在AP颗粒与基体之间的界面处设置接触损伤。利用有限元法，对含损伤颗粒夹杂模型进行计算，通过对比数值仿真结果，研究了固含量及界面损伤对复合固体推进剂细观损伤形貌及宏观力学性能的影响。结果表明，当固含量较低时，颗粒与基体之间界面损伤的聚合往往发生在少部分颗粒之间，随固含量增大，参与界面损伤聚合的颗粒逐渐增多，形成的宏观裂纹越来越明显；颗粒与基体之间的界面损伤，对复合固体推进剂非线性力学性能影响较大，不可忽略。

We report the synthesis of altitudinal molecular motors that contain functional groups in their rotor part. In an approach to achieve dynamic control over the properties of solid surfaces, a hydrophobic perfluorobutyl chain and a relatively hydrophilic cyano group were introduced to the rotor part o

We report the synthesis of altitudinal molecular motors that contain functional groups in their rotor part. In an approach to achieve dynamic control over the properties of solid surfaces, a hydrophobic perfluorobutyl chain and a relatively hydrophilic cyano group were introduced to the rotor part

Ammonium nitrate (AN) is extensively used in the area of fertilizers and explosives. It is present as the major component in most industrial explosives. Its use as an oxidizer in the area of propellants, however, is not as extensive as in explosive compositions or gas generators. With the growing demand for environmental friendly chlorine free propellants, many attempts have been made of late to investigate oxidizers producing innocuous combustion products. AN, unlike the widely used ammonium perchlorate, produces completely ecofriendly smokeless products. Besides, it is one of the cheapest and easily available compounds. However, its use in large rocket motors is restricted due to some of its adverse characteristics like hygroscopicity, near room temperature phase transformation involving a volume change, and low burning rate (BR) and energetics. The review is an attempt to consolidate the information available on the various issues pertaining to its use as a solidpropellant oxidizer. Detailed discussions on the aspects relating to phase modifications, decomposition chemistry, and BR and energetics of AN-based propellants, are presented. To make the review more comprehensive brief descriptions of the history, manufacture, safety, physical and chemical properties and various other applications of the salt are also included. Copyright 1999 Elsevier Science B.V.

The role of bonding agents on the aging characteristics is one of the important research topics.Aging program of the prepared propellant samples was conducting as follows:Five samples,two free of bonding agents,and three containing an aziridine based bonding agents (MAPO,HX-752,MAT4),four samples based on different bonding and curing agents all were aged at 70℃.The prepared bonding agent"MAT4"gave remarkable improvements and resulted in highly stable mechanical properties comparing with HX-752 or MAPO.The selected bonding agents family inhibited the rate of decomposition of the propellants during the aging periods and supported the propellant matrix against decomposition at the elevated temperatures.

Full Text Available A rapid, expedient and enantioselective method for the synthesis of beta-hydroxy amines and monosubstituted aziridines in up to 99% e.e., via asymmetric transfer hydrogenation of a-amino ketones and cyclisation through treatment with tosyl chloride and base, is described. (1R,2R-N-(para-toluenesulfonyl-1,2-ethylenediamine with formic acid has been utilised as a ligand for the Ruthenium (II catalysed enantioselective transfer hydrogenation of the ketones.The chiral 2-methyl aziridine, which is a potentially more efficient bonding agent for Rocket SolidPropellant has been successfully achieved.

transparent acrylic that allows cinematography of plasma flows and ignition events along the propellant bed. The chamber can withstand pressures up to ~13...formal averaging technique applied to the microscopic flow. These equations require a number of constitutive laws for closure including state equations

State-of-the-art composite propellants are based on solid particles (AP, Aluminium) in a polymeric HTPB based binder system. The usability of a propellant for a particular application is dependent on a large number of properties. These different properties sometimes result in contradictory requireme

In a prior effort, a thermal/fluid model of the interior of Penn State University's laboratory-scale Insulation Test Motor (ITM) was constructed to predict both the convective and radiative heat transfer to the interior walls of the ITM with a minimum of empiricism. These predictions were then compared to values of total and radiative heat flux measured in a previous series of ITM test firings to assess the capabilities and shortcomings of the chosen modeling approach. Though the calculated fluxes reasonably agreed with those measured during testing, this exercise revealed means of improving the fidelity of the model to, in the case of the thermal radiation, enable direct comparison of the measured and calculated fluxes and, for the total heat flux, compute a value indicative of the average measured condition. By replacing the P1-Approximation with the discrete ordinates (DO) model for the solution of the gray radiative transfer equation, the radiation intensity field in the optically thin region near the radiometer is accurately estimated, allowing the thermal radiation flux to be calculated on the heat-flux sensor itself, which was then compared directly to the measured values. Though the fully coupling the wall thermal response with the flow model was not attempted due to the excessive computational time required, a separate wall thermal response model was used to better estimate the average temperature of the graphite surfaces upstream of the heat flux gauges and improve the accuracy of both the total and radiative heat flux computations. The success of this modeling approach increases confidence in the ability of state-of-the-art thermal and fluid modeling to accurately predict SRM internal environments, offers corrections to older methods, and supplies a tool for further studies of the dynamics of SRM interiors.

The polyvinyl butyral (PVB)-based propellant films containing 78% CaCO3 were prepared by rolling process of light roller.The effects of PVB with different relative molecular mass and 1 1 kinds of liquid paraffins with different content of alcohols,acids and esters on the mechanical properties of propellant films were analyzed by means of statics mechanical test.The section morphology of the propellant films was observed by SEM.Results show that when PVB with relative molecular mass of 40 000 as an adhesive and liquid paraffin containing 5% alcohol as processing assistant are used,the internal structure of the propellant films is dense,the mechanical properties of propellant films are the best,with the low temperature tensile elongation of 2.97% and the high temperature tensile stress of 3.55 MPa.The rolling process of light roller is safe and stable for preparation of the PVB-based high solid content propellant when CaCO3 is substituted by RDX.The tensile stress and tensile elongation of propellant contai-ning RDX with particle size of 43.25μm are 45% and 17% higher than those of propellant containing RDX with par-ticle size of 21.02μm.%采用光辊压延工艺制备了含质量分数78％碳酸钙（CaCO3）的聚乙烯醇缩丁醛（PVB）基推进剂胶片。通过静态力学试验分析了不同相对分子质量PVB和11种液体石蜡对推进剂胶片力学性能的影响，用扫描电镜观察了推进剂胶片的剖片形貌。结果表明，以相对分子质量为40000的PVB作黏合剂、醇质量分数为5％的液体石蜡作工艺助剂时，推进剂胶片内部结构致密，力学性能最佳，低温延伸率为2．97％，高温拉伸强度为3．55 MPa。用黑索今（RDX）全部替代CaCO3时，可安全稳定地制备 PVB 基高固体含量推进剂。含43．25μm RDX 的推进剂比含21．02μm RDX的推进剂的拉伸强度及延伸率分别高45％和17％。

本文对航空座椅某型固体火箭发动机部装、总装及检测、试验、包装技术难点等进行了工艺分析；介绍了固体火箭发动机装配全过程工艺流程、检测、试验方法及注意事项等，对于同类及新型火箭发动机的装配制造过程具有良好的借鉴、推广应用意义。%Aiming at the difficulty of solidpropellant rocket engine of aviation seat to assembly, testing and packaging technology, the assembly, testing process and method for solidpropellant rocket engine were introduced. It can be regarded as reference with application for solidpropellant rocket engine assembly process.

Lift-off acoustic environments generated by the future Ares I launch vehicle are assessed by the NASA Marshall Space Flight Center (MSFC) acoustics team using several prediction tools. This acoustic environment is directly caused by the Ares I First Stage booster, powered by the five-segment Reusable Solid Rocket Motor (RSRMV). The RSRMV is a larger-thrust derivative design from the currently used Space Shuttle solid rocket motor, the Reusable Solid Rocket Motor (RSRM). Lift-off acoustics is an integral part of the composite launch vibration environment affecting the Ares launch vehicle and must be assessed to help generate hardware qualification levels and ensure structural integrity of the vehicle during launch and lift-off. Available prediction tools that use free field noise source spectrums as a starting point for generation of lift-off acoustic environments are described in the monograph NASA SP-8072: "Acoustic Loads Generated by the Propulsion System." This monograph uses a reference database for free field noise source spectrums which consist of subscale rocket motor firings, oriented in horizontal static configurations. The phrase "subscale" is appropriate, since the thrust levels of rockets in the reference database are orders of magnitude lower than the current design thrust for the Ares launch family. Thus, extrapolation is needed to extend the various reference curves to match Ares-scale acoustic levels. This extrapolation process yields a subsequent amount of uncertainty added upon the acoustic environment predictions. As the Ares launch vehicle design schedule progresses, it is important to take every opportunity to lower prediction uncertainty and subsequently increase prediction accuracy. Never before in NASA s history has plume acoustics been measured for large scale solid rocket motors. Approximately twice a year, the RSRM prime vendor, ATK Launch Systems, static fires an assembled RSRM motor in a horizontal configuration at their test facility

In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust measurements is pointed out. According to the measurement characteristics, a dynamic digital filter compensation method is presented. Combined the identification-modeling, dynamic compensation and simulation, the system's dynamic mathematic model is established. And then, a compensation digital filter is also designed. Thus, the dynamic response of the system is improved and the instantaneous thrust measurement can be implemented. The measurement results for the rocket motor show that the digital filter compensation is effective in the instantaneous thrust measurement.

Microthrusters are critical for the development of terrestrial micromissiles and nano air vehicles for reconnaissance, surveillance, and sensor emplacement. With the maturation of MEMS manufacturing technology, the physical components of the thrusters can be readily fabricated. The thruster type that is the most straightforward is chemical combustion of a propellant that is ignited by a heating element giving a single shot thrust. Arrays of MEMS manufactured thrusters can be ganged to give multiple firings. The basic model for such a system is a solid rocket motor. The desired elements for the propellant of a chemical thruster are high specific impulse (I sp), high temperature and pressure, and low molecular weight combustion gases. Since the combustion chamber of a microthruster is extremely small, the propellant material must be able to ignite, sustain and complete its burn inside the chamber. The propellant can be either a solid or a liquid. There are a large number of energetic materials available as candidates for a propellant for microthrusters. There has been no systematic evaluation of the available energetic materials as propellant candidates for microthrusters. This report summarizes computations done on a series of energetic materials to address their suitabilities as microthruster propellants.

Burning aluminized propellants eject reacting molten aluminum drops with a broad size distribution. Prior to this work, in situ measurement of the drop size statistics and other quantitative flow properties was complicated by the narrow depth-of-focus of microscopic videography. Here, digital in-line holography (DIH) is demonstrated for quantitative volumetric imaging of the propellant plume. For the first time, to the best of our knowledge, in-focus features, including burning surfaces, drop morphologies, and reaction zones, are automatically measured through a depth spanning many millimeters. By quantifying all drops within the line of sight, DIH provides an order of magnitude increase in the effective data rate compared to traditional imaging. This enables rapid quantification of the drop size distribution with limited experimental repetition.

SYNTHESIS OF LIQUID PROPELLANT Hydroxylammonium nitrate (HAN), prepared via the electrolysis of nitric acid, is commercially available as a high-purity...stack gases, and brine solution from the wet scrubber (82). 5 Applicability/Limitation Most types of solid, liquid, and gaseous organic wastes or

A methodology for the reliability analysis of a reusable solid rocket motor case is discussed. The analysis is based on probabilistic fracture mechanics and probability distribution for initial flaw sizes. The developed reliability analysis is used to select the structural design variables of the solid rocket motor case on the basis of minimum expected cost and specified reliability bounds during the projected design life of the case. Effects of failure prevention plans such as nondestructive inspection and the material erosion between missions are also considered in the developed procedure for selection of design variables. The reliability-based procedure can be modified to consider other similar structures of reusable space vehicle systems with different failure prevention plans.

The selection of a satisfactory numerical method for calculating the propagation of steep fronted shock life waveforms in a solid rocket motor combustion chamber is discussed. A number of different numerical schemes were evaluated by comparing the results obtained for three problems: the shock tube problems; the linear wave equation, and nonlinear wave propagation in a closed tube. The most promising method--a combination of the Lax-Wendroff, Hybrid and Artificial Compression techniques, was incorporated into an existing nonlinear instability program. The capability of the modified program to treat steep fronted wave instabilities in low smoke tactical motors was verified by solving a number of motor test cases with disturbance amplitudes as high as 80% of the mean pressure.

Full Text Available Numerical simulations are carried out for the internal flow field of a dual pulse solid rocket motor port to understand the flow behaviour. Three dimensional Reynolds Averaged Navier Stokes equations are solved alongwith shear stress transport turbulence model using commercial code. The combustion gas is assumed as a mixture of alumina and gases and single phase flow calculations are done with the thermo chemical properties provided for the mixture. The simulation captures all the essential features of the flow field. The flow accelerates through the pulse separation device (PSD port and high temperature and high velocity gas is seen to impinge the motor wall near the PSD port. The overall total pressure drop through motor port and through PSD is found to be moderate.Defence Science Journal, 2012, 62(6, pp.369-374, DOI:http://dx.doi.org/10.14429/dsj.62.1418

In this paper, we tackle the issue of the accurate simulation of evaporating and reactive polydisperse sprays strongly coupled to unsteady gaseous flows. In solid propulsion, aluminum particles are included in the propellant to improve the global performances but the distributed combustion of these droplets in the chamber is suspected to be a driving mechanism of hydrodynamic and acoustic instabilities. The faithful prediction of two-phase interactions is a determining step for future solid rocket motor optimization. When looking at saving computational ressources as required for industrial applications, performing reliable simulations of two-phase flow instabilities appears as a challenge for both modeling and scientific computing. The size polydispersity, which conditions the droplet dynamics, is a key parameter that has to be accounted for. For moderately dense sprays, a kinetic approach based on a statistical point of view is particularly appropriate. The spray is described by a number density function and its evolution follows a Williams-Boltzmann transport equation. To solve it, we use Eulerian Multi-Fluid methods, based on a continuous discretization of the size phase space into sections, which offer an accurate treatment of the polydispersion. The objective of this paper is threefold: first to derive a new Two Size Moment Multi-Fluid model that is able to tackle evaporating polydisperse sprays at low cost while accurately describing the main driving mechanisms, second to develop a dedicated evaporation scheme to treat simultaneously mass, moment and energy exchanges with the gas and between the sections. Finally, to design a time splitting operator strategy respecting both reactive two-phase flow physics and cost/accuracy ratio required for industrial computations. Using a research code, we provide 0D validations of the new scheme before assessing the splitting technique's ability on a reference two-phase flow acoustic case. Implemented in the industrial

propeller for ice operation ships. A typical propeller profile was created using MATLAB and modeled in SolidWorks using realistic material properties...for ice operation ships. A typical propeller profile was created using MATLAB and modeled in SolidWorks using realistic material properties. The...OPENPROP in MATLAB The program allows for the 3-D graphical propeller design created in MATLAB to be exported to CAD programs such as Rhino or SolidWorks

针对固体推进剂在压延塑化过程中温度过高而容易引起的燃爆问题,提出了远红外测温与喷雾降温。通过试验得出,当药料的温度超过设定值时,经喷雾降温后,药料的温度降低了3~5℃,有效提高了固体推进剂在压延塑化过程中的安全性。在此基础上,研究了在压延塑化过程中,影响药料温度变化的有关因素,得出了结论：随着工作辊转速、工作辊退水温度、压延塑化遍数和硝化棉含量的增加,药料的温度随之升高；随着硝化甘油含量的增加,药料的温度随之降低。%In view of the high temperature and being easy to cause combustion and explosion of solidpropellant in the process of rolling plasticizing,the method of far infrared temperature measurement and spray cooling is put forward,it is concluded through experiments that,when the temperature of solidpropellant exceeds the set value,then after the spray cooling,the temperature of solidpropellant reduces 3 - 5 ℃,it improves the security of solidpropellant in the process of rolling plasticizing effectively.On the basis of this,the effect factors of the solidpropellant temperature have been re-searched during the processing of rolling plasticizing.It is concluded that as the increase of work roll speed,the work roll returning water temperature,the number of rolling plasticizing,and the content of nitrocellulose(NC),the temperature of solidpropellant will increase,but as the increase of the content of nitroglycerine(NG),the temperature of solidpropellant will decrease instead.

Spontaneous motion of a solid/liquid composite induced by a chemical Marangoni effect, where an oil droplet attached to a solid soap is placed on a water phase, was investigated. The composite exhibits various characteristic motions, such as revolution (orbital motion) and translational motion. The results showed that the mode of this spontaneous motion switches with a change in the size of the solid scrap. The essential features of this mode-switching were reproduced by ordinary differential equations by considering nonlinear friction with proper symmetry.

总结了固体推进剂结构完整性分析的研究状况，综述了结构完整性分析的手段方法，并重点介绍了数值仿真在固体推进剂结构完整性分析中的应用。阐述了推进剂粘弹性力学特性和复杂的载荷环境对数值仿真分析造成的困难，以及目前利用数值仿真分析在固体推进剂结构完整性研究中取得的成就和未来的发展趋势。%The analysis of the structure integrity of solidpropellants is summarized .The methods for structure integrity analysis are reviewed, the applications of numerical simulation in the analysis of solidpropellant structure integrity are introduced especially .The difficulties of numerical simulation analysis caused by the propellant mechanical properties and the complexity of load environment are explained .The achievements and the development trend in the future of using numerical simulation analysis in the study of solidpropellant structure integrity are presented .

The dissipation characteristic of fatigue damage for composite solidpropellant was studied for evaluating the effects of vibration loading on solidpropellant. The mechanical and memory characteristic of propellant was analyzed. The characteristic of fatigue damage of propellant under vibration loading was dis-cussed. The critical stress cycle number and its change law were given. The cumulative damage model based on vibration dissipated energy was set up. The results can afford available help for calculating cumu-lative damage of solid charge under vibration of road transportation and shipbone.%为考察运输过程中振动载荷对固体推进剂的影响，进行了复合固体推进剂振动疲劳损伤耗散特性研究，分析了复合固体推进剂的力学特性和记忆特性，讨论了推进剂在振动载荷作用下的疲劳损伤特征，给出了临界应力循环数及变化规律，建立了基于振动耗散能的累积损伤模型，可以计算发动机固体装药在公路运输和舰载环境下的振动累积损伤。

The recertification of the Space Shuttle Solid Rocket Booster (SRB) and Solid Rocket Motor (SRM) has included an extensive rewriting of the Failure Mode and Effects Analysis (FMEA) and Critical Items List (CIL). The evolution of the groundrules and methodology used in the analysis is discussed and compared to standard FMEA techniques. Especially highlighted are aspects of the FMEA/CIL which are unique to the analysis of an SRM. The criticality category definitions are presented and the rationale for assigning criticality is presented. The various data required by the CIL and contribution of this data to the retention rationale is also presented. As an example, the FMEA and CIL for the SRM nozzle assembly is discussed in detail. This highlights some of the difficulties associated with the analysis of a system with the unique mission requirements of the Space Shuttle.

One of the most significant barriers encountered to the space application of MEMS technology is its lack of reliability and flight heritage in space environments. In this study a MEMS solidpropellant thruster array was selected for the verification test of MEMS technology in space. The function and performance of MEMS solid thruster have been previously verified by laboratory-level research in universities. To ensure the successful operation of the MEMS thruster module before flight demonstration on-orbit, launch and on-orbit environment tests were performed at the qualification level. In the launch test, sine burst, and random vibration loads were applied to the MEMS thruster module. The thermal vacuum tests were carried out for the on-orbit environment test. As a result of the launch vibration test and on-orbit environment test, the variations of the characteristics were less than 0.7%, and all the functional requirements were successfully verified after the vibration tests. The tests successfully verified the manufacturing process because the thruster module showed stable normal function before the ignition. The test result outputs will be helpful in establishing MEMS fabrication guidelines for space applications.

为了研究硝胺推进剂（挤压成型和浇注成型）和复合固体推进剂的断裂韧性，文章采用非接触类型的视频引伸仪对固体推进剂CT试样进行断裂测试，测得厚度为38 mm的3种试样断裂参数K F 分别为：0．7546、0．0812、0．3368 MPa· m1/2，并且建立推进剂材料的失效评估图。发现挤压成型的硝胺推进剂断裂韧性比浇注成型硝胺推进剂和复合推进剂高，CT试样预估的失效载荷与试验结果能较好地吻合。%To study the fracture toughness of nitramine (in extruded and slurry cast conditions ) and composite solidpropellants , a non-contacting type video extensometer was applied to test solidpropellant CT specimens .The fracture parameters KF for 3 samples with a thickness of 38mm were 0.7546,0.0812 and 0.3368MPa· m1/2, respectively.The failure assessment diagrams of propellant material were generated .Fracture toughness of the extruded nitramine propellant was found to be higher than those of slurry cast nitramine propellant and composite propellant .Failure load estimates of CT specimens were found to be in good agreement with test results .

Full Text Available The thermo-structural response of solid rocket motor nozzles is widely investigated in the design of modern rockets, and many factors related to the material properties have been considered. However, little work has been done to evaluate the effects of structure gaps on the generation of flame leaks. In this paper, a numerical simulation was performed by the finite element method to study the thermo-structural response of a typical nozzle with consideration of the structure gap. Initial boundary conditions for thermo-structural simulation were defined by a quasi-1D model, and then coupled simulations of different gap size matching modes were conducted. It was found that frictional interface treatment could efficiently reduce the stress level. Based on the defined flame leak criteria, gap size optimization was carried out, and the best gap matching mode was determined for designing the nozzle. Testing experiment indicated that the simulation results from the proposed method agreed well with the experimental results. It is believed that the simulation method is effective for investigating thermo-structural responses, as well as designing proper gaps for solid rocket motor nozzles.

Asbestos free solidmotor internal insulation samples were recently tested at the MSFC Hyperthermal Arc Jet Facility. Objectives of the test were to gather data for solid rocket motor analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Active instrumentation included use of total calorimeters, in-depth thermocouples, and a surface pyrometer for in-situ surface temperature measurement. Post-test sample forensics involved determination of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

Full Text Available A computer code has been developed for internal ballistic performance evaluation of solid rocket motors, using minimum distance function (MDF approach for prediction of geometry evolution. This method can handle any complex geometry without the need to define different geometrical shapes and their evolution as used in several existing analytical geometry evolution-based methodologies. The code is validated with both experimental results published in literature, as well as for solid rocket motors of tactical and strategic missiles and a very good match is obtained with static test results. The output of the code gives p-t (pressure-time curve as well as the detailed parameters of the flow along the axial direction, and geometries in the form of mesh file, which can be further used as input to codes for CFD analysis.Defence Science Journal, Vol. 65, No. 3, May 2015, pp.181-188, DOI: http://dx.doi.org/10.14429/dsj.65.8304

A capability to couple NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This paper summarizes the efforts in applying the installed coupling software to demonstrate/investigate fluid-structure interaction (FSI) between pressure wave and flexible inhibitor inside reusable solid rocket motor (RSRM). First a unified governing equation for both fluid and structure is presented, then an Eulerian-Lagrangian framework is described to satisfy the interfacial continuity requirements. The features of fluid solver, Loci/CHEM and structural solver, CoBi, are discussed before the coupling methodology of the solvers is described. The simulation uses production level CFD LES turbulence model with a grid resolution of 80 million cells. The flexible inhibitor is modeled with full 3D shell elements. Verifications against analytical solutions of structural model under steady uniform pressure condition and under dynamic condition of modal analysis show excellent agreements in terms of displacement distribution and eigen modal frequencies. The preliminary coupled result shows that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor.

Full Text Available An experimental investigation on the ignition of metallised propellants (APIHTPB/AI has been carried out 10 determine the ignition delay, minimum ignition energy and corresponding heat flux,threshold heat flux for ignition and minimum ignition temperature, Ignition experiments were conductedusing a shock tube under convectiveheating conditions similar to those prevailingin a rocket motor. Heat flux at propellant location was measured by thin film heat flux gauge and also calculated from a ribbon thermocouple output under similar test conditions. The igntion delay was measured as the time lag between the arrival of hot gas at the propellant and the light emission due to actual ignition of the propellant. The experimental results indicate that the ignition delay characteristics are independent of pressure. The minimum energy for ignition obtained for the propellant is 1100J/m2 corresponding to the heat flux range of 80·120 WIcm2 for a gas velocity of 110 mls. The threshold heat flux required to ignite the propellant was 40 W/cm2 at a velocity of 110 mls. Heat flux corresponding to minimum ignition energy and the threshold heat flux increase with gas velocity. The threshold ignition temperature of the propellant was found to be 600 ± 20 K.

Multi-stage launch vehicles with solid rocket motors (SRMs) face design optimization challenges, especially when the mission scope changes frequently. Significant performance benefits can be realized if the solid rocket motors are optimized to the changing requirements. While SRMs represent a fixed performance at launch, rapid design iterations enable flexibility at design time, yielding significant performance gains. The streamlining and integration of SRM design and analysis can be achieved with improved analysis tools. While powerful and versatile, the Solid Performance Program (SPP) is not conducive to rapid design iteration. Performing a design iteration with SPP and a trajectory solver is a labor intensive process. To enable a better workflow, SPP, the Program to Optimize Simulated Trajectories (POST), and the interfaces between them have been improved and automated, and a graphical user interface (GUI) has been developed. The GUI enables real-time visual feedback of grain and nozzle design inputs, enforces parameter dependencies, removes redundancies, and simplifies manipulation of SPP and POST's numerous options. Automating the analysis also simplifies batch analyses and trade studies. Finally, the GUI provides post-processing, visualization, and comparison of results. Wrapping legacy high-fidelity analysis codes with modern software provides the improved interface necessary to enable rapid coupled SRM ballistics and vehicle trajectory analysis. Low cost trade studies demonstrate the sensitivities of flight performance metrics to propulsion characteristics. Incorporating high fidelity analysis from SPP into vehicle design reduces performance margins and improves reliability. By flying an SRM designed with the same assumptions as the rest of the vehicle, accurate comparisons can be made between competing architectures. In summary, this flexible workflow is a critical component to designing a versatile launch vehicle model that can accommodate a volatile

利用铅柱压缩量、铅壔扩孔值和冲击波超压与冲量等参数定量研究了推进剂残余装药对战斗部爆炸毁伤效果的影响,并探讨了影响战斗部装药/推进剂体系铅柱压缩量的因素.试验结果表明,四种推进剂的存在对战斗部装药/推进剂体系的铅柱压缩量均有不同程度的增益作用.HTPB四组元推进剂对战斗部装药/推进剂体系的铅柱压缩量和铅壔扩孔值的贡献量要高于HTPB三组元推进剂( HTPB-3),其铅壔扩孔值的增益量分别为81.4％和54％.比例距离范围为3.5～12.5 m·kg-1/3时,HTPB-3对1 kg战斗部装药PBXN/0.6 kg HTPB-3体系的超压TNT当量和冲量TNT当量分别贡献了18.7％和19.7％.%The influence of solidpropellant on warhead damage effect was studied by these quantitative parameters including lead cylinder compression value,lead block volume,shock wave overpressure and impulse. Influencing factors that effect lead cylinder compression value of warhead charge/propellant system was also discussed. The results show that four kinds of propellants all could enhance lead cylinder compression value of warhead charge/propellant system in different degrees. Contribution to lead cylinder compression value of warhead charge/ propellant system by four groups HTPB propellant is more than that by three groups HTPB propellant ( HTPB-3). Increased percentage of lead block volume is 81.4% and 54% ,respectively. When scaled distance ranges from 3.5 m· kg-1/3 to 12.5 m ? Kg-1/3,HTPB-3 contributes to shock wave overpressure TNT equivalence and impulse TNT equivalence of 1 kg PBXN/0.6 kg HTPB-3 system by an increase of 18.7% and 19.7%,respectively.

Full Text Available A novel block polymer has been synthesised from caprolactone using hydroxy terminated polybutadiene as ring opening initiator. Usefulness of this polymer as propellant binder has been studied by generating data on physico-chemical properties of the polymer. The polymer exhibited high miscibility with nitrate ester and high solid loading capability. Preliminary data generated on typical propellant formulation indicated higher performance as compared to composite propellant.

In this paper ，viscoelastic property of solid rocket propellant is described by fractional derivative model．Frequency response function and impulse response function of one dimensional pole is deduced by means of Laplace transformation and its inverse transformation．The proportions of pole part and branchcut part in the response are analyzed．The transitional and steady response of the pole is calculated under harmonic excitation and the steady acceleration response is compared with expeimental results．%本文用三参数分数阶导数模型描述了固体推进剂的粘弹特性，用Laplace变换和反变换求解一维杆件的频响函数和脉冲响应函数，分析了极点部分和截断部分在响应中所占比例，计算了在简谐激励下杆件的瞬态和稳态响应，并就加速度稳态响应部分与试验做了对比。

Firstly,a simulation research was carried on,and the temperature characteristics changing with time under different power between the two electrodes was obtained.Then the temperature of arc plasma center was used as a boundary conditions for solidpropellant ignition,and solidpropellant ignition delay time under different power was calculated.Finally,come to a conclusion that the solidpropellant ignition delay time decreases with the increase of electrode power,when the electrode power increase from 50 W to 500 W,solidpropellant ignition delay time decreases from 8.82 ms to 0.99 ms,and with the increase of power,the growth of ignition delay time is slowing down.%首先对电弧进行了仿真研究，得到了不同电弧功率下电弧温度时间特性，然后利用该温度特性，作为固体推进剂点火的边界条件，采用数值差分方法，计算了固体推进剂的点火延迟时间；结果表明：固体推进剂的点火延迟时间随着电弧功率的增大而减小，当电弧功率从50 W 增加到500 W 时，固体推进剂的点火延迟时间从8．82 ms 减小到0．99 ms，并且随着电弧功率的升高，点火延迟时间减小的速度减慢。

For those case-bonded casting solid rocket motors ( SRMs) with large outside/inside radius ratio of the grain, high thermal strain will be generated in the propellant grain subjected to thermal loading, which significantly limit the performance of SRM. Pressure cure is an effective method to reduce the thermal strain in the grain. In this paper, the theory of pressure cure was analyzed and the relationship between the desired pressure and the parameters of the SRM was deduced. Then, a finite element method ( FEM) of pressure cure named two step method was proposed. Theoretical calculation and FEM simulation were carried out on a tube motor with four different material cases. The recommended pressure of pressure cure of four cases were given out.%对于药柱外/内径比(m数)很大的贴壁浇注式固体火箭发动机,在固化降温后,推进剂药柱内会产生显著的热应变,这严重限制了发动机的进一步高性能化. 加压固化是一种降低推进剂药柱内热应变的有效方法. 文中分析了加压固化的原理,推导出了加压固化所需压强与发动机参数之间的关系式,提出了一种两步分析法的加压固化有限元分析方法. 针对4种不同壳体材料的圆管发动机,进行了加压固化理论计算与有限元仿真分析,给出了4种壳体加压固化时的推荐压强.

Metal particles are widely used in space engineering to increase specific impulse and to supress acoustic instability of intra-champber processes. A numerical analysis of the internal injection-driven turbulent gas-particle flows is performed to improve the current understanding and modeling capabilities of the complex flow characteristics in the combustion chambers of solid rocket motors (SRMs) in presence of forced pressure oscillations. The two-phase flow is simulated with a combined Eulerian-Lagrangian approach. The Reynolds-averaged Navier-Stokes equations and transport equations of k - ε model are solved numerically for the gas. The particulate phase is simulated through a Lagrangian deterministic and stochastic tracking models to provide particle trajectories and particle concentration. The results obtained highlight the crucial significance of the particle dispersion in turbulent flowfield and high potential of statistical methods. Strong coupling between acoustic oscillations, vortical motion, turbulent fluctuations and particle dynamics is observed.

This study reports on modelling of an autosynchronous electric motor stator, operating at transient state. The developed model, of the modal type, includes around 20 nodes. The simulations showed that hot spots are localized on the winding heads and led to the choice of a solid-liquid phase change cooling system. The comparison between simulation and experiment permitted the identification of unknown parameters. The model gives a good accuracy during steady-state and in the rising temperature phase. The modelling of the phase change cooling is realized by the addition of two nodes. The sensitivity analysis to PCM properties shows that the hot spot temperature decreases with increasing conductivities, inertia and latent heat of melting of the PCM and with decreasing melting temperature. Gallium (metal melting at 30{sup o}C) is the best PCM for the cooling of hot spots and P116 paraffin is the best non-metallic PCM. (author)

During the developmental static fire program for NASAs Reusable Solid Rocket Motor-V (RSRMV), an anomalous erosion condition appeared on the nozzle Carbon Cloth Phenolic nose ring that had not been observed in the space shuttle RSRM program. There were regions of augmented erosion located on the bottom of the forward nose ring (FNR) that measured nine tenths of an inch deeper than the surrounding material. Estimates of heating conditions for the RSRMV nozzle based on limited char and erosion data indicate that the total heat loading into the FNR, for the new five segment motor, is about 40-50% higher than the baseline shuttle RSRM nozzle FNR. Fault tree analysis of the augmented erosion condition has lead to a focus on a thermomechanical response of the material that is outside the existing experience base of shuttle CCP materials for this application. This paper provides a sensitivity study of the CCP material thermo-structural response subject to the design constraints and heating conditions unique to the RSRMV Forward Nose Ring application. Modeling techniques are based on 1-D thermal and porous media calculations where in-depth interlaminar loading conditions are calculated and compared to known capabilities at elevated temperatures. Parameters such as heat rate, in-depth pressures and temperature, degree of char, associated with initiation of the mechanical removal process are quantified and compared to a baseline thermo-chemical material removal mode. Conclusions regarding postulated material loss mechanisms are offered.

This simulation involved a two-dimensional axisymmetric model of a full motor initial grain of the Reusable Solid Rocket Motor (RSRM) of the Space Transportation System (STS). It was conducted with CFD (computational fluid dynamics) commercial code FLUENT. This analysis was performed to: a) maintain continuity with most related previous analyses, b) serve as a non-vectored baseline for any three-dimensional vectored nozzles, c) provide a relatively simple application and checkout for various CFD solution schemes, grid sensitivity studies, turbulence modeling and heat transfer, and d) calculate nozzle convective heat transfer coefficients. The accuracy of the present results and the selection of the numerical schemes and turbulence models were based on matching the rocket ballistic predictions of mass flow rate, head end pressure, vacuum thrust and specific impulse, and measured chamber pressure drop. Matching these ballistic predictions was found to be good. This study was limited to convective heat transfer and the results compared favorably with existing theory. On the other hand, qualitative comparison with backed-out data of the ratio of the convective heat transfer coefficient to the specific heat at constant pressure was made in a relative manner. This backed-out data was devised to match nozzle erosion that was a result of heat transfer (convective, radiative and conductive), chemical (transpirating), and mechanical (shear and particle impingement forces) effects combined.

An assessment of the reliability of solid-state motor controllers for nuclear power plants is made. Available data on failure-rate and failure-mode data for solid-state motor controllers based on industrial operating experience is meager; the data are augmented by data on other solid-state power electronic devices that are shown to have components similar to those found in solid-state motor controllers. In addition to large nonnuclear solid-state adjustable-speed motor drives, the reliability of nuclear plant inverter systems and high-voltage solid-state dc transmission-line converters is assessed. Licensee Event Report analyses from several sources, the open literature, and personal communications are used to determine the realiability of solid-state devices typical of those expected to be used in nuclear power plants in terms of failures per hour.

目前通用的大型商业软件都仅提供线性黏弹性本构模型,无法胜任固体推进剂材料的大变形分析.基于Abaqus提供的用户材料子程序接口VUMAT (vectorized user defined material subroutine),选用适应性强的Swanson非线性黏弹性本构模型进行二次开发.本构模型采用Fortran语言编写,可在Abaqus增量求解过程中作为子程序调用.通过标准犬骨单轴拉伸算例,验证子程序的有效性.所开发子程序考虑几何和材料双重非线性,能应用于大型固体火箭发动机药柱结构完整性分析,弥补Abaqus仅含线性黏弹性本构模型的不足.详述材料子程序开发流程,可为用户扩充Abaqus 的材料模型提供参考.%The linear viscoelastic constitutive model,which in most popular FEM (finite element method) software,is unsuitable for solidpropellant with large deformation.Based on the subroutine VUMAT (vectorized user defined material subroutine),a nonlinear viscoelastic constitutive model developed by Swanson was implemented into Abaqus.The constitutive model was coded using Fortran,and can be called as a subroutine.The validify of the subroutine had been proven through the standard uniaxial tensile model.The subroutine can be used for the analysis of large solid rocket motor.The shortage of Abaqus which only has linear viscoelastic constitutive model is remedied.Some useful technology,including the detail process of implementing a material constitutive model,was presented.It can be referred for extending the material constitutive model in Abaqus.

National Aeronautics and Space Administration — A small Mars (or other celestial body) ascent vehicle is unlikely to achieve the necessary propellant fraction required to achieve orbit. Scaling down of liquid...

Full Text Available An advanced and intensive computational solution development is integrated with an asymptotic technique, to examine the impact of the combustion surface morphology on the generated rotational flow field in a solid rocket chamber with wide ranges of forcing frequencies. The simulated rectangular chamber is closed at one end and is open at the aft end. The upper and lower walls are permeable to allow steady and unsteady injected air to generate internal flow mimicking the flow field of the combustion gases in real rocket chamber. The frequencies of the unsteady injected flow are chosen to be very close or away from the resonance frequencies of the adapted chamber. The current study accounts for a wide range of wave numbers that reflect the complexity of real burning processes. Detailed derivation for Navier-Stokes equations at the four boundaries of the chamber is introduced in the current study. Qualitative comparison is performed with recent experimental work carried out on a two-inch hybrid rocket motor using a mixture of polyethylene and aluminum powder. The higher the percentage of aluminum powder in the mixture, the more the corrugations of the combustion surface. This trend is almost similar to the computational and analytical results of a simulated solid rocket chamber.

In order to improve the rotary motion of surface assembled light-driven molecular motors, tetra-acid-functionalized motors were bound to an amine-coated quartz surface without prior activation of the acid groups. In contrast to earlier bipodal motors, the tetravalent motor showed no significant reduction in the rotation speed when attached to a surface.

In order to improve the rotary motion of surface assembled light-driven molecular motors, tetra-acid-functionalized motors were bound to an amine-coated quartz surface without prior activation of the acid groups. In contrast to earlier bipodal motors, the tetravalent motor showed no significant redu

A computational fluid dynamics (CFD) analysis has been performed on the aft slot region of the Titan 4 Solid Rocket Motor Upgrade (SRMU). This analysis was performed in conjunction with MSFC structural modeling of the propellant grain to determine if the flow field induced stresses would adversely alter the propellant geometry to the extent of causing motor failure. The results of the coupled CFD/stress analysis have shown that there is a continual increase of flow field resistance at the aft slot due to the aft segment propellant grain being progressively moved radially toward the centerline of the motor port. This 'bootstrapping' effect between grain radial movement and internal flow resistance is conducive to causing a rapid motor failure.

The lead-free double base propellants, renewable propellants based on thermoplastic elastomer binder (TPE), green composite propellants were reviewed. The characteristics of various types of green propellants and the difficult problems in the development of propellants were summarized, and the trends in the development of green propellants were pointed out, such as nano-technology, high efficiency loading technology and complex technology of non-lead catalysts, synthesizing and applying technology of energetic thermoplastic elastomer, purifying technology of HNF, synthesizing technology of new energetic oxidant.etc.%综述了无铅双基系推进剂、可再生TPE推进剂和绿色复合推进剂的研制现状,总结了各类绿色推进剂的特点和发展过程中的技术难题,指出了绿色推进剂的一些技术发展方向,如非铅催化剂的纳米化技术、高效负载技术和复合技术,含能热塑性弹性体的合成及应用技术,硝仿肼(HNF)提纯技术以及新型高能氧化剂合成技术等.

This study seeks to determine the similarities in plume radiation between reduced and full-scale solid rocket models in ground test conditions through investigation of flow and radiation for a series of scale ratios ranging from 0.1 to 1. The radiative transfer equation (RTE) considering gas and particle radiation in a non-uniform plume has been adopted and solved by the finite volume method (FVM) to compute the three dimensional, spectral and directional radiation of a plume in the infrared waveband 2–6μm. Conditions at wavelengths 2.7μm and 4.3μm are discussed in detail, and ratios of plume radiation for reduced-scale through full-scale models are examined. This work shows that, with increasing scale ratio of a computed rocket motor, area of the high-temperature core increases as a 2 power function of the scale ratio, and the radiation intensity of the plume increases with 2–2.5 power of the scale ratio. The infrared radiation of plume gases shows a strong spectral dependency, while that of Al2O3 particles shows spectral continuity of gray media. Spectral radiation intensity of a computed solid rocket plume’s high temperature core increases sig-nificantly in peak radiation spectra of plume gases CO and CO2. Al2O3 particles are the major radi-ation component in a rocket plume. There is good similarity between contours of plume spectral radiance from different scale models of computed rockets, and there are two peak spectra of radi-ation intensity at wavebands 2.7–3.0μm and 4.2–4.6μm. Directed radiation intensity of the entire plume volume will rise with increasing elevation angle.

The paper deals with tip-modified propellers and the methods which, over a period of two decades, have been applied to develop such propellers. The development is driven by the urge to increase the efficiency of propellers and can be seen as analogous to fitting end plates and winglets to aircraft...... wings. The literature on four different designs is reviewed: the end-plate propeller; the two-sided, shifted end-plate propeller; the tip-fin propeller; and the bladelet propeller. The conclusion is that it is indeed possible to design tip-modified propellers that, relative to an optimum conventional...

The injection-flow issuing from a porous medium in the cold-flow simulation of internal port flows in solid rocket motors is characterized by a spatial instability termed pseudoturbulence that produces a rather non-uniform (lumpy) injection-velocity profile. The objective of this study is to investigate the interaction between the injection- and the developing axial-flows. The findings show that this interaction generally weakens the lumpy injection profile and affects the subsequent development of the axial flow. The injection profile is found to depend on the material characteristics, and the ensuing pseudoturbulence is a function of the injection velocity, the axial position and the distance from the porous wall. The flow transition (from laminar to turbulent) of the axial-flow is accelerated in flows emerging from smaller pores primarily due to the higher pseudoturbulence produced by the smaller pores in comparison to that associated with larger pores. In flows with rather uniform injection-flow profiles (weak or no pseudoturbulence), the axial and transverse velocity components in the porous duct are found to satisfy the sine/cosine analytical solutions derived from inviscid assumptions. The transition results from the present study are compared with previous results from surveyed literature, and detailed flow development measurements are presented in terms of the blowing fraction, and characterizing Reynolds numbers.

在不同温度和应变率下,对双基固体推进剂试件进行了单轴拉伸试验.提出了一种针对双基推进剂屈服值的判断方法,应用一元回归数理统计方法对双基推进剂力学性能数据进行了分析.结果表明,双基推进剂力学性能与温度和应变率具有明显的相关性,在233.15 K下近似为脆性材料,而在288.15 K和323.15 K下呈现明显的塑性流动.针对双基推进剂不同受载情况,提出了与推进剂屈服值相关的强度准则,以便为双基推进剂药柱设计及结构完整性分析奠定基础.%The uniaxial tensile tests about double-base propellant samples were carried out under different environmental temperature and strain rate. A judgment method for double-base propellant yield value was put forward,and the mechanical properties of the double-base propellant were analyzed by a regression mathematical statistics method. The results show that the mechanical properties of the double-base propellant is obviously dependent with temperature and strain rate, which approximatly appears as brittle material under 233. 15 K,and displays obvious plastic flow under 288. 15 K and 323. 15 K. According to different loading conditions of double-base propellant,the strength criterion that is relevant to propellant yield value was proposed,which lays the foundation for design and the structural integrity analysis of double-base propellant grain.

Repeated exposure of a propellant to liquid nitrogen causes thermal stress gradients within the material resulting in cracking and particle size reduction. This process is termed cryocycling. The authors conducted a feasibility study, combining experiments on both inert and live propellants with three modeling approaches. These models provided optimized cycle times, predicted ultimate particle size, and allowed crack behavior to be explored. Process safety evaluations conducted separately indicated that cryocycling does not increase the sensitivity of the propellants examined. The results of this study suggest that cryocycling is a promising technology for the demilitarization of tactical rocket motors.

Aiming at the problems concerning the composite solidpropellant with large deformation in the matrix and interface debonding around the filled particles, an improved Mori-Tanaka method was developed. At the same time, in order to verify its effectiveness, a simulation method for the interface debonding of the composite solidpropellant consists of random filled particles was also put forward. Finally, a computational case of one kind of composite solidpropellant was investigated. It is demonstrated that the resuits from these two methods are close and the improved Mori-Tanaka method is effective; the computational time can be reduced dramatically and the computational effectiveness can be increased greatly by using the improved Mori-Tanaka method.%针对复合固体推进剂基体的大变形和界面的脱粘问题,对Mori-Tanaka法进行了改进.同时,为验证该方法的有效性,针对推进剂颗粒随机填充的特点,提出了一种含非线性界面脱粘的数值方法.最后以某推进剂配方为算例,对2种方法的计算结果进行了比较.结果表明,2种方法结果接近,从而验证了改进的Mori-Tanaka法的有效性,且改进的Mori-Tanaka法计算量小,极大地提高了计算效率.

@@ For technological reasons many high-performance solid rocket motors are made from segmented propellant grains with non-uniform port geometry. In this paper parametric studies have been carried out to examine the geometric dependence of transient flow features in solid rockets with non-uniform ports. Numerical computations have been carried out in an inert simulator of solidpropellant rocket motor with the aid of a standard k-ω turbulence model. It was seen that the damping of the temperature fluctuation is faster in solid rocket with convergent port than with divergent port geometry. We inferred that the damping of the flow fluctuations using the port geometry is a meaningful objective for the suppression and control of the instability and/or pressure/thrust oscillations during the starting transient of solid rockets.

Hydrogen chloride (HCl) and aluminum oxide (Al2O3) are major exhaust products of solid rocket motors (SRM). Samples of calcination-produced alumina were exposed to continuously flowing mixtures of gaseous HCl/H2O in nitrogen. Transient sorption rates, as well as maximum sorptive capacities, were found to be largely controlled by specific surface area for samples of alpha, theta, and gamma alumina. Sorption rates for small samples were characterized linearly with an empirical relationship that accounted for specific area and logarithmic time. Chemisorption occurred on all aluminas studied and appeared to form from the sorption of about a 2/5 HCl-to-H2O mole ratio. The chemisorbed phase was predominantly water soluble, yielding chloride/aluminum III ion mole ratios of about 3.3/1 suggestive of dissolved surface chlorides and/or oxychlorides. Isopiestic experiments in hydrochloric acid indicated that dissolution of alumina led to an increase in water-vapor pressure. Dissolution in aqueous SRM acid aerosol droplets, therefore, might be expected to promote evaporation.

Interest in small unmanned aerial vehicles has increased dramatically in recent years. Hybrid vehicles which allow forward flight as a fixed wing aircraft and a true vertical landing capability have always had applications. Management of the available energy and noise associated with electric propeller propulsion systems presents many challenges. NASA Langley has developed the Greased Lightning 10 (GL-10) vertical takeoff, unmanned aerial vehicle with ten individual motors and propellers. All are used for propulsion during takeoff and contribute to acoustic noise pollution which is an identified nuisance to the surrounding users. A propeller test capability was developed to gain an understanding of how the noise can be reduced while meeting minimum thrust requirements. The designed propeller test stand allowed for various commercially available propellers to be tested for potential direct replacement of the current GL-10 propellers and also supported testing of a newly designed propeller provided by the Georgia Institute of Technology. Results from the test program provided insight as to which factors affect the noise as well as performance characteristics. The outcome of the research effort showed that the current GL-10 propeller still represents the best choice of all the candidate propellers tested.

A program of formulation and process development for ANB-3652 motorpropellant was conducted to validate design concepts and screen critical propellant composition and process parameters. Design experiments resulted in the selection of a less active grade of ferric oxide to provide better burning rate control, the establishment of AP fluidization conditions that minimized the adverse effects of particle attrition, and the selection of a higher mix temperature to improve mechanical properties. It is shown that the propellant can be formulated with AP and aluminum powder from various producers. An extended duration pilot plant run demonstrated stable equipment operation and excellent reproducibility of propellant properties. A similar program of formulation and process optimization culminating in large batch scaleup was conducted for ANB-3672 igniter propellant. The results for both ANB-3652 and ANB 37672 confirmed that their processing characteristics are compatible with full-scale production.

Full Text Available The ability to predict with some accuracy a given solid rocket motor’s performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size. The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor’s internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor’s internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale.

We report that liquids perform self-propelled motion when they are placed in contact with hot surfaces with asymmetric (ratchetlike) topology. The pumping effect is observed when the liquid is in the Leidenfrost regime (the film-boiling regime), for many liquids and over a wide temperature range. We propose that liquid motion is driven by a viscous force exerted by vapor flow between the solid and the liquid.

In the invention, the speeds of both propellers in a counterrotating aircraft propeller pair are measured. Each speed is compared, using a feedback loop, with a demanded speed and, if actual speed does not equal demanded speed for either propeller, pitch of the proper propeller is changed in order to attain the demanded speed. A proportional/integral controller is used in the feedback loop. Further, phase of the propellers is measured and, if the phase does not equal a demanded phase, the speed of one propeller is changed, by changing pitch, until the proper phase is attained.

It is very important to accurately predict the gas pressure, gas and solid temperature, as well as the amount of O-ring erosion inside the space shuttle Reusable Solid Rocket Motor (RSRM) joints in the event of a leak path. The scenarios considered are typically hot combustion gas rapid pressurization events of small volumes through narrow and restricted flow paths. The ideal method for this prediction is a transient three-dimensional computational fluid dynamics (CFD) simulation with a computational domain including both combustion gas and surrounding solid regions. However, this has not yet been demonstrated to be economical for this application due to the enormous amount of CPU time and memory resulting from the relatively long fill time as well as the large pressure and temperature rising rate. Consequently, all CFD applications in RSRM joints so far are steady-state simulations with solid regions being excluded from the computational domain by assuming either a constant wall temperature or no heat transfer between the hot combustion gas and cool solid walls.

To research the decrement of solidpropellant relaxation modulus, the single cell model is established based on the components of a solidpropellant.According to the solidpropellant relaxation modulus which obtained under 5% strain load, the binder relaxation modulus is inversed by the FEM and homogenization theory.On this basis, the relaxation modulus of the solidpropellant are studied when the solidpropellant under different strain loads, and the drawing process of the solidpropellant is computed. The variation of relaxation modulus in the drawing process of the solidpropellant is obtained based on Bo-ltzmann superposition principle and the generalized Hooke law.The results show that relaxation modulus of the composite solidpropellant is dependent on its strain loads, and the decrement of relaxation modulus is dependent on the loading history in the drawing process.%为研究复合固体推进剂松弛模量的衰减，依据复合固体推进剂组分建立单胞体模型。根据复合固体推进剂5％定应变下的松弛模量，利用有限元计算方法和均匀化理论对基体材料的松弛模量进行参数反演。在此基础上，研究了复合固体推进剂单胞体模型在不同定应变水平下的松弛模量，并计算了单胞体模型的拉伸过程。根据 Boltzmann 叠加原理及广义胡克定律计算了单胞体模型拉伸过程中松弛模量的变化。结果表明，复合固体推进剂松弛模量有明显的应变相关性，并且在拉伸过程中模量的衰减与载荷历程相关。

A series of cold flow heat transfer tests was conducted with a 7.5-percent scale model of the Space Shuttle Rocket Motor (SRM) to measure the heat transfer coefficients in the separated flow region around the nose of the submerged nozzle. Modifications were made to an existing 7.5 percent scale model of the internal geometry of the aft end of the SRM, including the gimballed nozzle in order to accomplish the measurements. The model nozzle nose was fitted with a stainless steel shell with numerous thermocouples welded to the backside of the thin wall. A transient 'thin skin' experimental technique was used to measure the local heat transfer coefficients. The effects of Reynolds number, nozzle gimbal angle, and model location were correlated with a Stanton number versus Reynolds number correlation which may be used to determine the convective heating rates for the full scale Space Shuttle Solid Rocket Motor nozzle.

The thermal decomposition of ammonium perchlorate (AP) and ammonium-perchlorate-based composite propellants is studied using the simultaneous thermogravimetric modulated beam mass spectrometry (STMBMS) technique. The main objective of the present work is to evaluate whether the STMBMS can provide new data on these materials that will have sufficient detail on the reaction mechanisms and associated reaction kinetics to permit creation of a detailed model of the thermal decomposition process. Such a model is a necessary ingredient to engineering models of ignition and slow-cookoff for these AP-based composite propellants. Results show that the decomposition of pure AP is controlled by two processes. One occurs at lower temperatures (240 to 270 C), produces mainly H{sub 2}O, O{sub 2}, Cl{sub 2}, N{sub 2}O and HCl, and is shown to occur in the solid phase within the AP particles. 200{micro} diameter AP particles undergo 25% decomposition in the solid phase, whereas 20{micro} diameter AP particles undergo only 13% decomposition. The second process is dissociative sublimation of AP to NH{sub 3} + HClO{sub 4} followed by the decomposition of, and reaction between, these two products in the gas phase. The dissociative sublimation process occurs over the entire temperature range of AP decomposition, but only becomes dominant at temperatures above those for the solid-phase decomposition. AP-based composite propellants are used extensively in both small tactical rocket motors and large strategic rocket systems.

Static tests of horizontally restrained rocket motors at the ATK facility in Promontory UT, USA result in the deposition of entrained soil and fuel combustion products, referred to as Test Fire Soil (TFS), over areas as large as 30-50 mile (80-130 km) and at distances up to 10-12 miles (16-20 km) from the test site. Chloride is the main combustion product generated from the ammonium perchlorate-aluminum based composite propellant. Deposition sampling/characterization and a 6-month field corrosivity study using mild steel coupons were conducted in conjunction with the February 25th 2010 FSM-17 static test. The TFS deposition rates at the three study sites ranged from 1 to 5 g/min/m. TFS contained significantly more chloride than the surface soil collected from the test site. The TFS collected during two subsequent tests had similarly elevated chloride, suggesting that the results obtained in this study are applicable to other tests assuming that the rocket fuel composition remains similar. The field-deployed coupons exposed to the TFS had higher corrosion rates (3.6-5.0 mpy) than paired non-exposed coupons (1.6-1.8 mpy). Corrosion rates for all coupons decreased over time, but coupons exposed to the TFS always had a higher rate than the non-exposed. Differences in corrosion rates between the three study sites were also observed, with sites receiving more TFS deposition having higher corrosion rates.

This illustration is a cutaway of the solid rocket booster (SRB) sections with callouts. The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solidpropellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment. The boosters are designed to survive water impact at almost 60 miles per hour, maintain flotation with minimal damage, and preclude corrosion of the hardware exposed to the harsh seawater environment. Under the project management of the Marshall Space Flight Center, the SRB's are assembled and refurbished by the United Space Boosters. The SRM's are provided by the Morton Thiokol Corporation.

A new, thermal-flow simulation code, called SFLOW. has been developed to model the gas dynamics, heat transfer, as well as O-ring and flow path erosion inside the space shuttle solid rocket motor joints by combining SINDA/Glo, a commercial thermal analyzer. and SHARPO, a general-purpose CFD code developed at Thiokol Propulsion. SHARP was modified so that friction, heat transfer, mass addition, as well as minor losses in one-dimensional flow can be taken into account. The pressure, temperature and velocity of the combustion gas in the leak paths are calculated in SHARP by solving the time-dependent Navier-Stokes equations while the heat conduction in the solid is modeled by SINDA/G. The two codes are coupled by the heat flux at the solid-gas interface. A few test cases are presented and the results from SFLOW agree very well with the exact solutions or experimental data. These cases include Fanno flow where friction is important, Rayleigh flow where heat transfer between gas and solid is important, flow with mass addition due to the erosion of the solid wall, a transient volume venting process, as well as some transient one-dimensional flows with analytical solutions. In addition, SFLOW is applied to model the RSRM nozzle joint 4 subscale hot-flow tests and the predicted pressures, temperatures (both gas and solid), and O-ring erosions agree well with the experimental data. It was also found that the heat transfer between gas and solid has a major effect on the pressures and temperatures of the fill bottles in the RSRM nozzle joint 4 configuration No. 8 test.

Report presents a theory which gives a complete picture and an exact quantitative analysis of the whole phenomenon of the working of blade screws, but also unites in a continuous whole the entire scale of states of work conceivable for a blade screw. Chapter 1 is devoted to the establishment of the system of fundamental equations relating to the blade screw. Chapter 2 contains the general discussion of the 16 states of work which may establish themselves for a blade screw. The existence of the vortex ring state and the whirling phenomenon are established. All the fundamental functions which enter the blade-screw theory are submitted to a general analytical discussion. The general outline of the curve of the specific function is examined. Two limited cases of the work of the screw, the screw with a zero constructive pitch and the screw with an infinite constructive pitch, are pointed out. Chapter 3 is devoted to the study of the propulsive screw or propeller. (author)

使用固体姿控小火箭是实现动能拦截器快响应和高精度姿态控制的最佳方案之一。针对一种新型动能拦截器姿控小火箭布局，提出了点火组合混合搜索算法。描述了动能拦截器姿控小火箭的配置方案，分析了弹体自旋需求。设计了一种结合目标排序法和区间搜索法的点火组合混合搜索算法：当可用小火箭个数较少时，采用目标排序法；当可用小火箭个数较多时，采用区间搜索法。指令力矩近似仿真结果及姿态控制数值仿真结果表明：该算法能够有效地近似指令力矩，实现快速高精度的姿态跟踪。%Using solid-propellant nozzles is one of the best schemes for kinetic interceptor to realize the fast response and high precision of attitude control.A mixed searching algorithm for ignition combination was presented for a novel attitude control solid-propellant nozzle in kinetic interceptor.Firstly,the configuration of solid-propellant nozzles was described and spin requirements of the kinetic interceptor were analyzed.Then the mixed searching algorithm was designed by a combination of sorting method and interval searching method.Sorting method is selected when the number of available nozzles is small and interval searching method is chosen on the contrary.Results of instruction torque approximation simulation and attitude control numerical simulation suggest that the algorithm can track the instruction torque effectively and achieve attitude tracking quickly and with a high precision.

Full Text Available The search for a smokeless propellant has encouraged scientists and engineers to look for a chlorine-free oxidizer as a substitute for AP (ammonium perchlorate. Endeavors seemed to come to an end when ADN (ammonium dinitramide appeared in the West in the early 1990s. Although some drawbacks soon became apparent by that time, the foremost obstacle for its use in rocket-motors came from the patent originally applied for in the United States in 1990. Furthermore, environmental concerns have also increased during these two decades. Ammonium perchlorate is believed to cause thyroid cancer by contaminating soil and water. In addition, AP produces hydrogen chloride during burning which can cause acid rain and ozone layer depletion. Unlike AP, ADN stands for both smokeless and green propellant. Since then, much progress has been made in its development in synthesis, re-shaping, microencapsulation and solidpropellant. The high solubility of ADN in water has also allowed its application as liquid monopropellant. Tests have revealed Isp (specific impulse superior to that normally observed with hydrazine, one of the most harmful and hazardous liquid propellants. With constraints of use, along with the patent near to expiry, scientists and engineers are rushing to complete developments and patents until then.

Full Text Available The ability to understand and predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. A numerical model incorporating pertinent elements, such as a representative transient, frequency-dependent combustion response to pressure wave activity above the burning propellant surface, is applied to the investigation of scale effects (motor size, i.e., grain length and internal port diameter on influencing instability-related behaviour in a cylindrical-grain motor. The results of this investigation reveal that the motor’s size has a significant influence on transient pressure wave magnitude and structure, and on the appearance and magnitude of an associated base pressure rise.

Real-time radiography was successfully applied to the Titan-IV Solid Rocket Motor Upgrade (SRMU) static firing test QM-2 conducted February 22, 1993 at Phillips Laboratory, Edwards AFB, CA. The real-time video data obtained in this test gave the first incontrovertible evidence that the molten slag pool is low (less than 5 to 6 inches in depth referenced to the bottom of the aft dome cavity) before T + 55 seconds, builds fairly linearly from this point in time reaching a quasi-equilibrium depth of 16 to 17 inches at about T + 97 seconds, which is well below the top of the vectored nozzle, and maintains that level until T + 125 near the end motor burn. From T + 125 seconds to motor burn-out at T + 140 seconds the slag pool builds to a maximum depth of about 20 to 21 inches, still well below the top of the nozzle. The molten slag pool was observed to interact with motions of the vectored nozzle, and exhibit slosh and wave mode oscillations. A few slag ejection events were also observed.

The team performed a literature review, conducted residual stress measurements, performed failure analysis, and demonstrated a solid state additive manufacturing repair technique on samples removed from a scrapped propeller hub. The team evaluated multiple options for hub repair that included existing metal buildup technologies that the Federal Aviation Administration (FAA) has already embraced, such as cold spray, high velocity oxy-fuel deposition (HVOF), and plasma spray. In addition the team helped Piedmont Propulsion Systems, LLC (PPS) evaluate three potential solutions that could be deployed at different stages in the life cycle of aluminum alloy hubs, in addition to the conventional spray coating method for repair. For new hubs, a machining practice to prevent fretting with the steel drive shaft was recommended. For hubs that were refurbished with some material remaining above the minimal material condition (MMC), a silver interface applied by an electromagnetic pulse additive manufacturing method was recommended. For hubs that were at or below the MMC, a solid state additive manufacturing technique using ultrasonic welding (UW) of thin layers of 7075 aluminum to the hub interface was recommended. A cladding demonstration using the UW technique achieved mechanical bonding of the layers showing promise as a viable repair method.

Full Text Available Within the EMAD Research Center of the USV have been developed since 1996 a number of solutions in heliothermic actuators and motors area made on the basis of solid heating medium. The solutions relates to rotary actuators and motors as well as linear actuators with limited movement. Are presented the primary constructive solutions with the peculiarities and finally, the general conclusion regarding the experimentation and practical implementation.

A devescovinid flagellate from termites exhibits rapid gliding movements only when in close contact with other cells or with a substrate. Locomotion is powered not by the cell's own flagella nor by its remarkable rotary axostyle, but by the flagella of thousands of rod bacteria which live on its surface. That the ectosymbiotic bacteria actually propel the protozoan was shown by the following: (a) the bacteria, which lie in specialized pockets of the host membrane, bear typical procaryotic flagella on their exposed surface; (b) gliding continues when the devescovinid's own flagella and rotary axostyle are inactivated; (c) agents which inhibit bacterial flagellar motility, but not the protozoan's motile systems, stop gliding movements; (d) isolated vesicles derived from the surface of the devescovinid rotate at speeds dependent on the number of rod bacteria still attached; (e) individual rod bacteria can move independently over the surface of compressed cells; and (f) wave propagation by the flagellar bundles of the ectosymbiotic bacteria is visualized directly by video-enhanced polarization microscopy. Proximity to solid boundaries may be required to align the flagellar bundles of adjacent bacteria in the same direction, and/or to increase their propulsive efficiency (wall effect). This motility-linked symbiosis resembles the association of locomotory spirochetes with the Australian termite flagellate Mixotricha (Cleveland, L. R., and A. V. Grimstone, 1964, Proc. R. Soc. Lond. B Biol. Sci., 159:668-686), except that in our case propulsion is provided by bacterial flagella themselves. Since bacterial flagella rotate, an additional novelty of this system is that the surface bearing the procaryotic rotary motors is turned by the eucaryotic rotary motor within.

Full Text Available Heat of decomposition of the double base propellant has been calculated from the knowledge of the composition and decomposition enthalpy of the ingredients. This was compared with the experimentally observed value of the propellant decomposition which suggested that condensed phase contribution is very marginal (one twentieth of the total calorimetric value in double base propellants unlike composite solidpropellants where condensed phase contribution is as high as one third. The condensed-phase in double base propellant was attributed to the nitration of the 2n-diphenylamine stabilizer in the propellant matrix.

initiators, JA2 19-Perf Hex propellants and Hagedorn-Plastic manufactured Nitrocellulose with Alkohol - German manufacturer. STAR-ATO goal is to develop... propellants in the U.S. Army’s small, medium and large caliber munitions are all nitrocellulose -based. As the Army drives continuous improvement in both...understanding the influence of nitrocellulose properties on propellant performance. Projectiles are getting heavier, ammunition is being exposed to

对复合固体推进剂及复合固体推进剂/衬层粘接试件微CT扫描后重构的图像进行了分析和统计,得到了其细观形貌特征及推进剂颗粒典型尺寸.对颗粒填充算法进行了扩展,并以最小代表体积元为基础构建了复合固体推进剂/衬层粘接体系二维细观数值模型.使用cohesive单元表征颗粒/基体和推进剂/衬层界面,分析了不同老化时间复合固体推进剂/衬层在单向拉伸作用下的脱粘过程.结果表明该数值模型及分析方法能从细观上准确地模拟复合固体推进剂/衬层的脱粘过程.为细观层面上研究复合固体推进剂/衬层脱粘提供了有效的数值模型和分析方法.%Statistical analysis of the reconstructed image of the composite solidpropellant/liner scanned by mi-cro-CT is conducted. It' s mesostructure and typical propellant particle size is attained. Particle packing algorithm is extended and 2D numerical model of the composite solidpropellant/liner bonding system is constructed based on representative volume element. Cohesive element is used to simulate the interface of particle/matrix and propellant/ liner. The action of the composite solidpropellant/liner debonding process of the mesostructure model under uniaxi-al tension during different aging time is analyzed. The result shows that the mesostructure numerical model can simulate the composite solidpropellant/liner debonding process. Effective numerical model and analytical methods for the meso level research on composite solidpropellant/liner debonding process is offered.

In an analysis of 30 families affected by spinal muscular atrophy (SMA) we have used the solid-phase minisequencing method to determine the ratio between the number of telomeric and centromeric copies of the survival motor neuron gene (SMN and cBCD541 respectively) on normal and SMA chromosomes...

Full Text Available To prepare the porous NC-based (nitrocellulose-based gun propellants, the batch foaming process of using supercritical CO2 as the physical blowing agent is used. The solubilities of CO2 in the single-base propellants and TEGDN (trimethyleneglycol dinitrate propellants are measured by the gravimetric method, and SEM (scanning electron microscope is used to observe the morphology of foamed propellants. The result shows that a large amount of CO2 could be dissolved in NC-based propellants. The experimental results also reveal that the energetic plasticizer TEGDN exerts an important influence on the pore structure. The triaxial tensile failure mechanism for solid-state nucleation is used to explain the nucleation of NC-based propellants in the solid state. Since some specific foaming behaviors of NC-based propellants can not be explained by the failure mechanism, a solid-state nucleation mechanism which revises the triaxial tensile failure mechanism is proposed and discussed.

Full Text Available Experimental and computer methods were developed for investigating the combustion phenomena in the propellants which burn in streams of hot gas flowing along the burn surfaces of the propellants. The experimental investigations allowed establishment of different dependencies for erosive burning. Computer solutions of the problem for double-base (DB propellants showed a good agreement with the experimental results. The suggested variant of modified theory considers the change of heat release in solids, the real burn surface roughness, the nonisothermality of boundary layer and the effect of gas mass blow from the propellant burn surface into the gas stream. This modified theory was used for studying burn laws at 30-1000 atm and up to gas stream sound velocities for different DB propellants. It was found that gas stream leads to splitting of the propellant burn laws, m = bp/sup v/. Pressure power (v, in this case depends on gas stream velocity (W, diameter of the propellant tube canal (d and gas stream temperature (T/sub w/. It is because of this that these burn laws were named partial burn laws. They have the form (m = bp/sup w(omega/ w,d,T/sub w/ -const. The dependencies w(omega = f(w,d,T/sub w/ were obtained by the modified theory. It was found that omega values mainly decrease when pressure increases beginning from ~200 to 400 atm and they can decrease up to w(omega = 0,1- 0,3. Similar results can be obtained for composite propellants.

Some key technologies solved during high-energy engine development were presented,including development of high energy solidpropellant and appropriate liner and the design of high volumetric grain,as well as correlative experiments,burning rate correlation,throat erosion behavior, the storage of solidpropellant, interface compatibility, safety, etc. Some important problems which need to be concerned in the next step were also discussed.%对高能发动机研制中解决的系列关键技术进行了总结,包括高能推进剂及配套衬层的研制、高装填药柱结构设计以及相关试验验证、燃速相关性、喉衬烧蚀特性、推进剂贮存性、界面相容性、安全性等,提出了还需继续关注和深入研究的问题.

Full Text Available Ignition of a composite aluminised propellant (AP-HTPB-Al in stagnant hot air is analysed theoretically, based on solid phase and gas phase theories. According to solid phase theory, ignition is due to reaction of the propellant in the solid phase at elevated temperatures. One-dimensional transient solid phase energy equation is solved to obtain the surface temperature profile of the propellant. By gas phase theory, an exothermic gas phase reaction, adjacent to the propellant surface, is considered responsible for the ignition. The changes in temperature and concentrations in the gas phase and the temperature profile below the propellant surface during the pre-ignition induction period are considered. Equations of energy and concentrations of reactants have been solved to obtain the species concentration and temperature profiles in the gas phase. An experimental investigation of the ignition of AP-HTPB-Al propellant is also carried out in a shock tube under end-mount conditions. Pressure and temperature ranges were 6-16 bar and 1500-3000 K, respectively. A comparison of the experimental data with predicted results shows that the ignition in an oxidizing atmosphere is by gas phase reaction, whereas in an inert atmosphere, solid phase reaction may be predominant.

Effects of metal (Ni, Cu, Al) and composite metal (NiB, NiCu, NiCuB) nanopowders on the thermal decomposition of ammonium perchlorate (AP) and composite solidpropellant ammonium perchlorate/hydroxyterminated polybutadiene (AP/HTPB) were studied by thermal analysis (DTA). The results show that metal and composite metal nanopowders all have good catalytic effects on the thermal decomposition of AP and AP/HTPB composite solidpropellant. The effects of metal nanopowders on the thermal decomposition of AP are less than those of the composite metal nanopowders. The effects of metal and composite metal nanopowders on the thermal decomposition of AP are different from those on the thermal decomposition of the AP/HTPB composite solidpropellant.

Based on the thermochemical erosion theory, a two-dimensional axisymmetric, coupled gas-solid-thermal numerical framework was established to predict the carbon-based nozzle erosion in solid rocket motors. Numerical simulations were carried out using the Wall Surface Reaction model of the commercial code FLUENT and the assumption whether the erosion process was chemical kinetics or diffusion controlled was not needed. The method was introduced to simulate the 70-1b BATES motor nozzle erosion and examine the effects of propellant composition, oxidizing species and chamber pressure. The calculated results agree well with experimental data. The erosion rate follows the trend exhibited by the heat flux distribution, and peaks slightly upstream of the throat. The erosion rate decreases with increasing aluminum content and increases almost linearly with chamber pressure. H20 is the dominant oxidizing species in dictating nozzle erosion.%为了准确预示固体火箭发动机碳基材料喷管的烧蚀率,依据热化学烧蚀理论,建立了喷管传热烧蚀的二维轴对称气-固-热耦合计算模型,计算通过FLUENT壁面化学反应模型完成,无需事先假设烧蚀控制机制。针对70-lb BATES发动机喷管进行了烧蚀计算,研究了推进剂配方、氧化性组分、燃烧室压强对喷管烧蚀的影响。结果表明：烧蚀率计算值与试验测试值吻合较好;烧蚀率分布遵循喷管内壁热流密度分布规律,在喉部上游入口处达到峰值;烧蚀率随推进剂Al含量增加而降低,随燃烧室压强升高而近似正比例增大;H2O是决定烧蚀的主要氧化性组分。

Based on the model of one-dimensional steady-state reaction gas flow, the correction factors of burning rate related to double base propellant, modified double base propellant and composite solidpropellant were summed up, the application scope of theoretical combustion model was broadened. Using Visual C++ and Microsoft Access for the development tools, the software of solidpropellant burning rate prediction(SPRS) was completed by the structural parameters of chemical bonds. The software was based on system of Windows XP, user-friendly, easy to use ,and with the functions of data updating and information querying. The burning rates and pressure indexs could been calculated when the chemical compositions of the propellant (formula) and the pressures were given. The compositions of the propellant(formula) could be adjusted by giving the burning rates and pressure indexs too. It was of great significance in development of cycle-shortening and cost-saving of solidpropellant.%在一维气相稳态反应流模型的基础上,总结了适用于双基推进剂、改性双基推进剂、复合固体推进剂燃速预估的修正因子,拓宽了燃烧理论模型的适用范围.采用Visual C++和Microsoft Access为开发工具,完成了基于组分化学键结构参数的固体推进剂燃速预估软件(SPRS)编制.该软件基于Windows XP系统,界面友好,使用方便,具有数据更新和信息查询功能.用户不仅能根据推进剂的化学组成(配方)和给定压强计算燃速、压力指数等参数,还可根据给定的燃速和压力指数等调整推进剂配方组成,对缩短固体推进剂研制周期和节约研制成本具有重要意义.

通过对固体推进剂药柱进行初始气孔率热老化试验研究,探讨了用累积损伤方法来预测药柱贮存寿命,提出了一种准确、经济、方便地提前预测药柱贮存寿命的新方法和新技术.%A method for predicting the storage life of solidpropellant grain by cumulative damage was developed based on heat ageing experiment on initial void content rate of propellant grain,and its application was discussed.A new method which can accurately,economically and conveniently predict the storage life of solid propillant grain was prosoed.

Intense high frequency (25-38 kHz) tone bursts have been observed in acoustic tests of a scale model of a general aviation propeller. The amplitude of the tone burst is approximately equal to the amplitude of the propeller noise signature. The conditions necessary for the production of these tone bursts are described. The experiments indicate that the origin of these bursts is a periodic flow oscillation on the suction surface of the propeller blade tips which may be due to the interaction between an oscillating shock wave and a laminar boundary layer.

The aim of this study was to examine if a propeller thoracodorsal artery perforator (TAP) flap can be used for breast reconstruction. Fifteen women were reconstructed using a propeller TAP flap, an implant, and an ADM. Preoperative colour Doppler ultrasonography was used for patient selection...... major complications needing additional surgery. One flap was lost due to a vascular problem. Breast reconstruction can be performed by a propeller TAP flap without cutting the descending branch of the thoracodorsal vessels. However, the authors would recommend that a small cuff of muscle is left around...

A numerical model for calculating performance of solidpropellant air-turbo-rocket at design point was set up,and a relationship expression of fuel-air ratio of combustion based on compressor pressure ratio,tubine intet total temperature and turbine expansion ratio was put forward.And a relationship between the turbine expansion ratio and the bypass ratio was given.The effects of compressor pressure ratio, turbine inlet temperature,bypass ratio/turbine expansion ratio and flight Mach number on the specific thrust and specific impulse were analyzed quantitatively.%建立了固体推进剂吸气式涡轮火箭发动机的设计状态数值模型,提出了基于压气机增压比、涡轮前温度和涡轮落压比关系的燃烧室燃气与空气配比表达式,以及涡轮落压比和发动机涵道比的匹配关系.定量分析了压气机增压比、涡轮进口燃气总温、涵道比/涡轮落压比和飞行马赫数对固体推进剂吸气式涡轮火箭发动机的单位推力和比冲的影响.

Burning rate and specific impulse of a solidpropellant are extremely important parameters in a design of a solid rocket motor. In this study, the relations between the burning rate and the amount of energy contained in HMX-CMDB propellants wherein the amount of energy is varied by adding HMX (High Melting Point Explosive). The following results are obtained. The final flame temperature is getting higher when the amount of energy is increased by adding HMX into a double-base propellant. The higher the final flame temperature is, the lower the burning rate is. Dark zone temperature, as a physical property, is lowered when the containing amount of energy is increased by adding HMX into the double-base propellant. This is because that, when weight fraction of HMX is increased, reaction heat at burning surface decreases, and the reaction in fizz zone is getting slower. The higher the dark zone temperature is, the higher the burning rate is. 20 refs., 11 figs., 1 tab.

Results of a series of noise measurements that were made under controlled conditions during the static firing of two Nike solidpropellant rocket motors are presented. The usefulness of these motors as sources for general spacecraft noise testing was assessed, and the noise expected in the cargo bay of the orbiter was reproduced. Brief descriptions of the Nike motor, the general procedures utilized for the noise tests, and representative noise data including overall sound pressure levels, one third octave band spectra, and octave band spectra were reviewed. Data are presented on two motors of different ages in order to show the similarity between noise measurements made on motors having different loading dates. The measured noise from these tests is then compared to that estimated for the space shuttle orbiter cargo bay.

In this paper, a design method for increasing performance of the marine propellers including the WCT propeller is suggested. It is described to maximize the performance of the propeller by adjusting expanded areas of the propeller blade. Results show that efficiency can be increased up to over 2% through the suggested design method.

The Advanced Solid Rocket Motor (ASRM), currently in its design and development phase, will become the higher reliability/performance next-generation booster for the Space Shuttle. The ASRM design improves safety through the elimination of 229 potential leak paths, thereby eliminating a total of 312 causes of failure. The ASRM will also allow an additional 12,000 lb of payload to be carried by the Shuttle Orbiter. Construction of the ASRM will promote competition through the use of a government owned/contractor operated manufacturing facility. Continuous-mix propellant grain manufacture will be used to enhance safety, quality, and efficiency.

The Advanced Solid Rocket Motor (ASRM), currently in its design and development phase, will become the higher reliability/performance next-generation booster for the Space Shuttle. The ASRM design improves safety through the elimination of 229 potential leak paths, thereby eliminating a total of 312 causes of failure. The ASRM will also allow an additional 12,000 lb of payload to be carried by the Shuttle Orbiter. Construction of the ASRM will promote competition through the use of a government owned/contractor operated manufacturing facility. Continuous-mix propellant grain manufacture will be used to enhance safety, quality, and efficiency.

This book deals with flows over propellers operating behind ships, and the hydrodynamic forces and moments which the propeller generates on the shaft and on the ship hull.The first part of the text is devoted to fundamentals of the flow about hydrofoil sections (with and without cavitation......) and about wings. It then treats propellers in uniform flow, first via advanced actuator disc modelling, and then using lifting-line theory. Pragmatic guidance is given for design and evaluation of performance, including the use of computer modelling.The second part covers the development of unsteady forces...... arising from operation in non-uniform hull wakes. First, by a number of simplifications, various aspects of the problem are dealt with separately until the full problem of a non-cavitating, wide-bladed propeller in a wake is treated by a new and completely developed theory. Next, the complicated problem...

Research efforts are being conducted by the NASA Lewis Research Center to formulate and characterize the properties of Al/RP-1 and RP-1 gelled propellants for rocket propulsion systems. Twenty four different compositions of gelled fuels were formualted with 5 and 16 micron, atomized aluminum powder in RP-1. The total solids concentration in the propellant varied from 5 to 60 wt percent. Tests were conducted to evaluate the stability and rheological characteristics of the fuels. Physical separation of the solids occurred in fuels with less than 50 wt percent solids concentration. The rheological characteristics of the Al/RP-1 fuels varied with solids concentration. Both thixotropic and rheopectic gel behavior were observed. The unmetallized RP-1 gels, which were formulated by a different technique than the Al/RP-1 gels, were highly viscoelastic. A history of research efforts which were conducted to formulate and characterize the properties of metallized propellants for various applications is also given.

By way of measuring the acoustic emission (AE) signals of Hydroxyl-terminated polybutadiene (HTPB) propellant in condition of uniform speed, and combined with the scanning electron microscopy (SEM) fracture surface observation, the damage law of HTPB composite solidpropellant under thermomechanical loading was studied. The results show that the effects of thermomechanical loading on HTPB propellant are related to the time and can be divided into three different stages. In the first stage, thermal air aging dominates; in the second stage, interface damage is dominant; and in the third stage, thermal air aging is once again dominant.

The Leidenfrost phenomenon of liquid droplets levitating and dancing when placed upon a hot plate due to propulsion of evaporative vapor has been extended to many self-propelled circumstances. However, such self-propelled Leidenfrost devices commonly need a high temperature for evaporation and a structured solid substrate for directional movements. Here we observed a “cold Leidenfrost phenomenon” when placing a dry ice device on the surface of room temperature water, based on which we develop...

We report that liquids perform self-propelled motion when they are placed in contact with hot surfaces with asymmetric (ratchet-like) topology. The pumping effect is observed when the liquid is in the film-boiling regime, for many liquids and over a wide temperature range. We propose that liquid motion is driven by a viscous force exerted by vapor flow between the solid and the liquid.

It was realized as early as 1909 that a propeller in yaw develops a side force like that of a fin. In 1917, R. G. Harris expressed this force in terms of the torque coefficient for the unyawed propeller. Of several attempts to express the side force directly in terms of the shape of the blades, however, none has been completely satisfactory. An analysis that incorporates induction effects not adequately covered in previous work and that gives good agreement with experiment over a wide range of operating conditions is presented. The present analysis shows that the fin analogy may be extended to the form of the side-force expression and that the effective fin area may be taken as the projected side area of the propeller.

Self-propelled droplets are a special kind of self-propelled matter that are easily fabricated by standard microfluidic tools and locomote for a certain time without external sources of energy. The typical driving mechanism is a Marangoni flow due to gradients in the interfacial energy on the droplet interface. In this article we review the hydrodynamic prerequisites for self-sustained locomotion and present two examples to realize those conditions for emulsion droplets, i.e. droplets stabilized by a surfactant layer in a surrounding immiscible liquid. One possibility to achieve self-propelled motion relies on chemical reactions affecting the surface active properties of the surfactant molecules. The other relies on micellar solubilization of the droplet phase into the surrounding liquid phase. Remarkable cruising ranges can be achieved in both cases and the relative insensitivity to their own `exhausts' allows to additionally study collective phenomena.

Full Text Available A blend of 3-carene and cardanol in 70:30 weight proportion exhibits synergistic hypergolic ignition with red fuming nitric acid (RFNA as oxidizer. Attempts have been made to evaluate this new propellant by theoretical calculationof performance parameters and verification of the results by static firing of a 10 kg thrust rocket motor around 20 atmosphers of chamber pressure. At an oxidizer-to-fuel weight ratio (O/F of 3.34 (RFNA used had 21% N204 and 5% by weight of concentrated sulphuric acid as catalyst, the propellant produced a reasonably smooth pressure-time curve with an ignition delay of 35 milliseconds. The theoretical characteristic velocity value matched well with the experimental. No carbon residue was left in the rocket motor after firing. Specific impulse (theoretical of the propellant has been found to be 223.8 seconds at chamber pressure, 20 atmos and exist pressure, 1 atmos.

Foamed propellants are based on crystalline explosives bonded in energetic reaction polymers. Due to their porous structures they are distinguished by high burning rates. Energy content and material characteristics can be varied by using different energetic fillers, energetic polymers and porous structures. Foamed charges can be produced easily by the reaction injection moulding process. For the manufacturing of foamed propellants a semi-continuous remote controlled production plant in pilot scale was set up and a modified reaction injection moulding process was applied. (Abstract Copyright [2009], Wiley Periodicals, Inc.)

A dynamic marine propeller simulation system was developed,which is utilized for meeting the experimental requirement of theory research and engineering design of marine electric propulsion system.By applying an actual ship parameter and its accurate propeller J'～KT' and J'～Kp' curve data,functional experiments based on the simulation system were carried out.The experiment results showed that the system can correctly emulate the propeller characteristics,produce the dynamic and steady performances of the propeller under different navigation modes,and present actual load torque for electric propulsion motor.

In order to avert the accidents of solid rocket motor explosion in fire and preserve warship's life-force and combat effectiveness, a one-dimensional mathematical model about the cook-off of the solid rocket motor of the ship-based missiles was established in this paper. The effects of radiation heat transfer, convective heat transfer,heat conduct and chemical reaction were considered in the model. The temperature-time curves, temperature-space curves and self-ignite delay times of the motor were obtained by means of solving numerically. The results indicated that there were marked effects of the temperature and the temperature rise rate of the fire on the self-ignite delay times of the motor, there was good adiabatical effect of the insulation of the motor on the fire and the ignition of the propellant appears firstly on the out surface under fast cook-off.%为了能够有效避免火灾中导弹固体火箭发动机着火爆炸的灾难发生、保存舰艇的生命力和战斗力,建立了一种舰载导弹固体火箭发动机烤燃过程的一维传热数学模型.该数学模型考虑了辐射换热、对流换热、导热和化学反应源项的作用.利用有限差分方法,通过数值计算得到了发动机的温度时间分布曲线、温度空间分布曲线和着火延迟时间.结果表明,火焰温度和火焰温升速率对发动机的着火延迟时间有显著影响;发动机的绝热层对外界火灾有好的隔热作用;快速热烤下,推进剂的着火首先发生在外表面上.

Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

Several formulations of propellants based on RDX as an energetic solid ingredients and cellulose acetate (CA) as a binder were processed using either dioctyl pthalate(DOP) or tracetin(TA) as plastisizer and a small amount of nitrocellulose(NC). The Performance of these propellants was evaluated on the basis of closed vessel firing data. The vulnerability aspects of these formulations were compared with those of conventional picrite propellant, NQ on the basis of their ignition temperat...

Full Text Available This article presents a dynamic regime of exposure of the radiant flux on the sample of gun-cotton. Obtained time the ignition of gun-cotton in the heating conditions of increasing heat flux in the range from 0.2 W/cm2 to 22 W/cm2. A comparison of the delay times of the ignition when heated variable and constant heat flux.

Many high temperature bodies such as flame, in which chemical reactions are very complex, emit their own spectra. These emission spectra usually consist of the spectral lines, spectral bands and the continuous spectra. In some cases, the spectral lines gather together. It is very difficult to find the right single spectral line when the spectral line intensity method is used. To deal with this problem, the idea that the single spectral line intensity is replaced by the total intensity of many spectral lines to measure the temperature is mentioned. And the relative intensity method is also changed to deal with this idea. The measurement of the temperature distribution based on this improved method is successful, and the measurement results are compared with the results of the thermocouple method.

An object of this invention is to provide a composition of beryllium hydride and carboxy-terminated polybutadiene which is stable. Another object of this invention is to provide a method for inhibiting the reactivity of beryllium hydride toward carboxy-terminated polybutadiene. It was found that a small amount of lecithin inhibits the reaction of beryllium hydride with the acid groups in carboxy terminated polybutadiene.

The binders tested were polyurethanes made from two hydroxy-terminated polybutadienes, R-45 and Butarez HT, one hydroxy-terminated butadiene-acrylonitrile copolymer, Hycar 1300X 17, and a hydroxy-terminated prepolymer, Esterdiol 560, made from the dimerized fatty acid Empol 1010. The isocyanates used most extensively were isophorone diisocyanate (IPDI) and a polymeric diisocyanate, DDI. Stress relaxation was used to examine the chemical changes that took place in the binder when subjected to the sterilization temperatures. The thermal stability of the oxidizer, ammonium perchlorate (AP), was tested by thermogravimetry in the isothermal and nonisothermal modes. The effect of particle size, recrystallization, moisture content, and doping on the heat stability of AP could be evaluated by this method. The volatile degradation products, obtained when AP samples were aged at 135 C for prolonged periods, were analyzed by mass spectroscopy.

An evaluation of the NASA's Marshall Space Flight Center (MSFC) strategy to implement Total Quality Management (TQM) in the Advanced Solid Rocket Motor (ASRM) Project is presented. The evaluation of the implementation strategy reflected the Civil Service personnel perspective at the project level. The external and internal environments at MSFC were analyzed for their effects on the ASRM TQM strategy. Organizational forms, cultures, management systems, problem solving techniques, and training were assessed for their influence on the implementation strategy. The influence of ASRM's effort was assessed relative to its impact on mature projects as well as future projects at MSFC.

Full Text Available Using the dispersed metal in solidpropellants to increase the temperature of combustion products leads to such a problem as the specific impulse loss due to the incomplete combustion of metal particles in the exhaust products. A redistribution of metal loaded into the propellant grain is one of the methods to decrease the specific impulse loss. This paper reports on the ways to obtain the optimum metal particle disposition for the case-bounded propellant grain of tube cross-sectional type. Three different approaches to analyze the metal combustion efficiency are discussed. The influence of the dynamic nonequilibrium of two-phase flow on the optimum metal particles disposition in the propellant grain of tube cross-sectional type is investigated.

The Leidenfrost phenomenon of liquid droplets levitating and dancing when placed upon a hot plate due to propulsion of evaporative vapor has been extended to many self-propelled circumstances. However, such self-propelled Leidenfrost devices commonly need a high temperature for evaporation and a structured solid substrate for directional movements. Here we observed a “cold Leidenfrost phenomenon” when placing a dry ice device on the surface of room temperature water, based on which we developed a controllable self-propelled dry ice hovercraft. Due to the sublimated vapor, the hovercraft could float on water and move in a programmable manner through designed structures. As demonstrations, we showed that the hovercraft could be used as a cargo ship or a petroleum contamination collector without consuming external power. This phenomenon enables a novel way to utilize programmable self-propelled devices on top of room temperature water, holding great potential for applications in energy, chemical engineering and biology.

The combustion of composite solidpropellants was investigated and an available numerical model was improved for taking into account the change of pressure, when the process occurs in a confined environment, as inside a rocket. The pressure increase upon ignition is correctly described by the improved model for both sandwich and dispersed particles propellants. In the latter case, self-induced fluctuations in the pressure and in all other computed variables occur, as consequence of the periodic rise and depletion of oxidizer particles from the binder matrix. The comparison with the results of the constant pressure model shows a different fluctuating profile of gas velocity, with a possible second order effect induced by the pressure fluctuations.

Full Text Available Ammonium-nitrate-(AN- based composite propellants prepared with a hydroxyl-terminated polybutadiene (HTPB/polytetrahydrofuran (PTHF blend binder have unique thermal decomposition characteristics. In this study, the burning characteristics of AN/HTPB/PTHF propellants are investigated. The specific impulse and adiabatic flame temperature of an AN-based propellant theoretically increases with an increase in the proportion of PTHF in the HTPB/PTHF blend. With an AN/HTPB propellant, a solid residue is left on the burning surface of the propellant, and the shape of this residue is similar to that of the propellant. On the other hand, an AN/HTPB/PTHF propellant does not leave a solid residue. The burning rates of the AN/HTPB/PTHF propellant are not markedly different from those of the AN/HTPB propellant because some of the liquefied HTPB/PTHF binder cover the burning surface and impede decomposition and combustion. The burning rates of an AN/HTPB/PTHF propellant with a burning catalyst are higher than those of an AN/HTPB propellant supplemented with a catalyst. The beneficial effect of the blend binder on the burning characteristics is clarified upon the addition of a catalyst. The catalyst suppresses the negative influence of the liquefied binder that covers the burning surface. Thus, HTPB/PTHF blend binders are useful in improving the performance of AN-based propellants.

The burning rate characteristics of solid composite propellants can be modified via different methods. One of these is the application of different oxidizer particle sizes in the propellant. The effect of the use of fine and coarse particles of the oxidizer hydrazinium nitroformate (HNF) in hydroxyl

A two-dimensional finite element model for predicting fracture patterns obtained in high energy gas fracture experiments is presented. In these experiments, a mixture of propellants is used instead of explosives to fracture the rock surrounding the borehole. The propellant mixture is chosen to tailor the pressure pulse so that multiple fractures emanate from the borehole. The model allows the fracture pattern and pressure pulse to be calculated for different combinations of propellant mixture, in situ stress conditions, and rock properties. The model calculates the amount of gas generated by the burning propellants using a burn rate given by a power law in pressure. By assuming that the gas behaves as a perfect gas and that the flow down the fractures is isothermal, the loss of gas from the borehole due to flow down the cracks is accounted for. The flow of gas down the cracks is included in an approximate manner by assuming self-similar pressure profiles along the fractures. Numerical examples are presented and compared to three different full-scale experiments. Results show a good correlation with the experimental data over a wide variety of test parameters. 9 reference, 10 figures, 3 tables.

针对C/C喉衬喷管小推力长时间工作固体火箭发动机,分别开展了含铝、不含铝两种推进剂状态的地面试验。根据燃烧室压强及发动机推力测试曲线计算了喷管喉径的瞬变值,对比研究了喉衬的烧蚀、沉积过程,指出含铝推进剂发动机C/C喉衬先后经历初始沉积、沉积消融、持续烧蚀、烧蚀与沉积交替四个阶段,而推进剂不含铝时则没有明显的初始沉积与沉积消融。讨论了推进剂配方、燃烧室压强、喷管结构等因素对喉衬烧蚀、沉积的影响,并提出了相应的改善措施。%In order to study the erosion and deposition process of the carbon-carbon nozzle inserts in low-thrust,long-duration solid rocket applications,two motors using metallized and non-metallized propellants,respectively,were tested.The transient value of the throat diameter was obtained from the pressure and thrust measurements.It was indicated that for metallized propellant the throat experienced four stages as initial deposition,deposition melt,continuous erosion,alternately erosion and deposition,while no initial deposition and melt phenomena were observed in the non-metallized case.Some approaches to minimize throat variation were developed based upon the investigation of the effect of chamber pressure,propellant composition and nozzle configurations on the throat erosion and deposition.

Full Text Available Several formulations of propellants based on RDX as an energetic solid ingredients and cellulose acetate (CA as a binder were processed using either dioctyl pthalate(DOP or tracetin(TA as plastisizer and a small amount of nitrocellulose(NC. The Performance of these propellants was evaluated on the basis of closed vessel firing data. The vulnerability aspects of these formulations were compared with those of conventional picrite propellant, NQ on the basis of their ignition temperatures and sensitivity to friction and impact. Triacetin was found to be better plasticizer than DOP for CA binder. Some RDX/CA/TA/NC/-based propellants were found to have energy levels comparable with NQ propellant and had less sensitivity to heat, impact and friction, and therefore have the potential for being used as low-vulnerability ammunition propellants for gun applications.

There is a strategic interest in understanding how the propellant manufacturing process contributes to military capabilities outside the United States. The paper will discuss how system dynamics (SD) has been applied to rapidly assess the capabilities and vulnerabilities of a specific composite propellant production complex. These facilities produce a commonly used solidpropellant with military applications. The authors will explain how an SD model can be configured to match a specific production facility followed by a series of scenarios designed to analyze operational vulnerabilities. By using the simulation model to rapidly analyze operational risks, the analyst gains a better understanding of production complexities. There are several benefits of developing SD models to simulate chemical production. SD is an effective tool for characterizing complex problems, especially the production process where the cascading effect of outages quickly taxes common understanding. By programming expert knowledge into an SD application, these tools are transformed into a knowledge management resource that facilitates rapid learning without requiring years of experience in production operations. It also permits the analyst to rapidly respond to crisis situations and other time-sensitive missions. Most importantly, the quantitative understanding gained from applying the SD model lends itself to strategic analysis and planning.

There is a strategic interest in understanding how the propellant manufacturing process contributes to military capabilities outside the United States. The paper will discuss how system dynamics (SD) has been applied to rapidly assess the capabilities and vulnerabilities of a specific composite propellant production complex. These facilities produce a commonly used solidpropellant with military applications. The authors will explain how an SD model can be configured to match a specific production facility followed by a series of scenarios designed to analyze operational vulnerabilities. By using the simulation model to rapidly analyze operational risks, the analyst gains a better understanding of production complexities. There are several benefits of developing SD models to simulate chemical production. SD is an effective tool for characterizing complex problems, especially the production process where the cascading effect of outages quickly taxes common understanding. By programming expert knowledge into an SD application, these tools are transformed into a knowledge management resource that facilitates rapid learning without requiring years of experience in production operations. It also permits the analyst to rapidly respond to crisis situations and other time-sensitive missions. Most importantly, the quantitative understanding gained from applying the SD model lends itself to strategic analysis and planning.

An analysis of several atomic hydrogen launch vehicles was conducted. A discussion of the facilities and the technologies that would be needed for these vehicles is also presented. The Gross Liftoff Weights (GLOW) for two systems were estimated; their specific impulses (I{sub sp}) were 750 and 1500 lb{sub f}/s/lb{sub m}. The atomic hydrogen launch vehicles were also compared to the currently planned Advanced Launch System design concepts. Very significant GLOW reductions of 52 to 58 percent are possible over the Advanced Launch System designs. Applying atomic hydrogen propellants to upper stages was also considered. Very high I{sub sp} (greater than 750 lb{sub f}/s/lb{sub m}) is needed to enable a mass savings over advanced oxygen/hydrogen propulsion. Associated with the potential benefits of high I(sub sp) atomic hydrogen are several challenging problems. Very high magnetic fields are required to maintain the atomic hydrogen in a solid hydrogen matrix. The magnetic field strength was estimated to be 30 kilogauss (3 Tesla). Also the storage temperature of the propellant is 4 K. This very low temperature will require a large refrigeration facility for the launch vehicle. The design considerations for a very high recombination rate for the propellant are also discussed. A recombination rate of 210 cm/s is predicted for atomic hydrogen. This high recombination rate can produce very high acceleration for the launch vehicle. Unique insulation or segmentation to inhibit the propellant may be needed to reduce its recombination rate.

In solidpropellants, aluminum is widely used to improve the performance, however the condensed combustion products especially the large agglomerates generated from aluminum combustion significantly affect the combustion and internal flow inside the solid rocket motor. To clarify the properties of the condensed combustion products of aluminized propellants, a constant-pressure quench vessel was adopted to collect the combustion products. The morphology and chemical compositions of the collected products, were then studied by using scanning electron microscopy coupled with energy dispersive (SEM-EDS) method. Various structures have been observed in the condensed combustion products. Apart from the typical agglomerates or smoke oxide particles observed before, new structures including the smoke oxide clusters, irregular agglomerates and carbon-inclusions are discovered and investigated. Smoke oxide particles have the highest amount in the products. The highly dispersed oxide particle is spherical with very smooth surface and is on the order of 1-2 μm, but due to the high temperature and long residence time, these small particles will aggregate into smoke oxide clusters which are much larger than the initial particles. Three types of spherical agglomerates have been found. As the ambient gas temperature is much higher than the boiling point of Al2O3, the condensation layer inside which the aluminum drop is burning would evaporate quickly, which result in the fact that few "hollow agglomerates" has been found compared to "cap agglomerates" and "solid agglomerates". Irregular agglomerates usually larger than spherical agglomerates. The formation of irregular agglomerates likely happens by three stages: deformation of spherical aluminum drops; combination of particles with various shape; finally production of irregular agglomerates. EDS results show the ratio of O to Al on the surface of agglomerates is lower in comparison to smoke oxide particles. C and O account for

National Aeronautics and Space Administration — The Cryogenic Propellant Storage and Transfer project will demonstrate the capability to safely and efficiently store, transfer and measure cryogenic propellants,...

An analytical model of the steady-state burning of composite solidpropellants is presented. An improved burning rate model is achieved by incorporating an improved AP monopropellant model, a separate energy balance for the binder in which a portion of the diffusion flame is used to heat the binder, proper use of the binder regression rate in the model, and a model for the combustion of the energetic binder component of CMDB propellants. Also, an improved correlation and model of aluminum agglomeration is developed which properly describes compositional trends.

Based on three-dimension viscoelastic finite element method, by MSC/NASTRAN software system, the grain structural integrity of the motor under internal pressure load, under thermal load and under the combined action of the two load is analyzed and evaluatedrespectively. The results show that the grain structural integrity of the solid rocket motor meets the design requirement satisfaction.%基于三维粘弹性有限元模型,应用MSC/NASTRAN软件对某发动机分别在固化降温、燃气内压载荷条件下的装药结构完整性进行分析,并对该发动机在固化降温、燃气内压两种载荷联合作用下的装药结构完整性进行评估。结果表明,该发动机的装药结构完整性满足要求。

Full Text Available The article presents dynamic mechanical analysis (DMA for solid rocket propellants testing. Principles of operation and measured values are briefly described. The authors refer to the previous research of PTFE material and literature data providing information about proper experimental conditions and influence of measurement frequency, load amplitude, and heating rate on the results of DMA tests. The experimental results of solid double-base rocket propellant testing obtained on the N Netzsch DMA 242 device are presented. Mechanical properties such as the dynamic storage modulus E´, the dynamic loss modulus E˝ and tan(δ were measured within temperature range from (–120°C to (+90°C at the heating rate of 1 K/min. The test sample was subjected to a dual cantilever multi-frequency test. Special attention was paid to determination of the glass transition temperature of the tested propellant in reference to the NATO standardization agreement 4540 as well as influence of the measurement frequency on the glass transition.[b]Keywords[/b]: Dynamic mechanical analysis, solid rocket propellants, glass transition temperature

With the technology currently available for the manufacture of propellants, it is possible to control the variance of the total specific impulse obtained from the rocket boosters to within approximately five percent. Though at first inspection this may appear to be a reasonable amount of control, when it is considered that any uncertainty in the total kinetic energy delivered to the spacecraft translates into a design with less total usable payload, even this degree of uncertainty becomes unacceptable. There is strong motivation to control the variance in the specific impulse of the shuttle's solid boosters. Any small gains in the predictability and reliability of the booster would lead to a very substantial payoff in earth-to-orbit payload. The purpose of this study is to examine one aspect of the manufacture of solidpropellants, namely, the mixing process. The traditional approach of computational fluid mechanics is notoriously complex and time consuming. Certain simplifications are made, yet certain fundamental aspects of the mixing process are investigated as a whole. It is possible to consider a mixing process in a mathematical sense as an operator, F, which maps a domain back upon itself. An operator which demonstrates good mixing should be able to spread any subset of the domain completely and evenly throughout the whole domain by successive applications of the mixing operator, F. Two and three dimensional models are developed and graphical visualization two and three dimensional mixing processes are presented.

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

The three commonly used propellant systems - H2/O2, RP-1/O2, and solidpropellants - primarily radiate as molecular emitters, non-scattering small particles, and scattering larger particles, respectively. Present technology has accepted the uncoupling of the radiation analysis from that of the flowfield. This approximation becomes increasingly inaccurate as one considers plumes, interior rocket chambers, and nuclear rocket propulsion devices. This study will develop a hierarchy of methods which will address radiation/convection coupling in all of the aforementioned propulsion systems. The nature of the radiation/convection coupled problem is that the divergence of the radiative heat flux must be included in the energy equation and that the local, volume-averaged intensity of the radiation must be determined by a solution of the radiative transfer equation (RTE). The intensity is approximated by solving the RTE along several lines of sight (LOS) for each point in the flowfield. Such a procedure is extremely costly; therefore, further approximations are needed. Modified differential approximations are being developed for this purpose. It is not obvious which order of approximations are required for a given rocket motor analysis. Therefore, LOS calculations have been made for typical rocket motor operating conditions in order to select the type approximations required. The results of these radiation calculations, and the interpretation of these intensity predictions are presented herein.

Full Text Available Composite propellant suspensions consist of highly filled polymeric system wherein solid particles of different sizes and shapes are dispersed in a polymeric matrix. The rheological behaviour of a propellant suspension is characterised by viscoplasticity and shear rate and time dependant viscosity. The behaviour of composite propellant suspension has been studied under amplitude sweep test where tests were performed by continuously varying strain amplitude (strain in %, γ by keeping the frequency and temperature constant and results are plotted in terms of log γ (strain amplitude vs logGʹ and logGʺ (Storage modulus and loss modulus, respectively. It is clear from amplitude sweep test that dynamic moduli and complex viscosity show marked increase at critical strain amplitude after a plateau region, infering a shear thickening behaviour.

A method was developed for the design of propellers of minimum induced loss matched to an arbitrary operating point characterized by disc loading (thrust or power), air density, shaft speed, flight speed, and number of blades. A consistent procedure is outlined to predict the performance of these propellers under off design conditions, or to predict the performance of propellers of general geometry. The examples discussed include a man powered airplane, a hang glider with a 7.5 kW (10 hp) 8,000 rpm engine, and an airplane-like motorsoarer.

Rapidly spinning magnetic white dwarfs in symbiotic stars may pass through the propeller stage. It is believed that a magnetic propeller acts in two such stars CH Cyg and MWC 560. We review a diversity of manifestations of the propeller there. In these systems in a quiescent state the accretion onto a white dwarf from the strong enough wind of a companion star is suppressed by the magnetic field, and the hot component luminosity is low. Since the gas stored in the envelope eventually settles ...